A quantitative figure-of-merit approach for optimization of an unmanned Mars Sample Return mission
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Authors Preiss, Bruce Kenneth, 1964-
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A quantitative figure-of-merit approach for optimization of an unmanned Mars Sample Return mission
Preiss, Bruce Kenneth, M.S.
The University of Arizona, 1991
U MI 300 N. Zeeb Rd. Ann Arbor, MI 48106
A QUANTITATIVE FIGURE-OF-MERIT APPROACH
FOR OPTIMIZATION OF AN UNMANNED MARS SAMPLE RETURN MISSION
by Bruce Kenneth Preiss
A Thesis Submitted to the Faculty of the DEPARTMENT OF AEROSPACE AND MECHANICAL ENGINEERING In Partial Fulfillment of the Requirements For the Degree of MASTER OF SCIENCE WITH A MAJOR IN AEROSPACE ENGINEERING In the Graduate College THE UNIVERSITY OF ARIZONA
1991 STATEMENT BY AUTHOR
This thesis has been submitted in partial fulfillment of requirements for an advanced degree at The University of Arizona and is deposited in the University Library to be made available to borrowers under rules of the library.
Brief quotations from this thesis are allowable without special permission, provided that accurate acknowledgment of source is made. Requests for permission for extended quotation from or reproduction of this manuscript in whole or in part may be granted by the head of the major department or the Dean of the Graduate College when in his or her judgment the proposed use of the material is in the interests of scholarship. In all other instances, however, permission must be obtained from the author.
SIGNED: ^
APPROVAL BY THESIS DIRECTOR
This thesis has been approved on the date shown below:
r. U-^-—iQ-CjC- °( I
K.N.R. Ramohalli Date Professor of Aerospace Engineering Ill ACKNOWLEDGMENTS
I would like to express my deepest gratitude to:
Dr. Kumar Ramohalli for his guidance and extreme patience. Without his influence and extensive knowledge, this project could not have succeeded.
Tom Pan for his dedicated help in many areas. His tireless research and comprehensive computer graphics work enabled the project to be completed on time.
Mario Rascon for his propellant background advice and extensive computer support.
Heidi Ruffner, my wife, for her infinite understanding and resolute moral support. Her excellent proofreading skills and merciless editing talent helped keep the writing on track. I have yet to figure out how she found the time to help me and still maintain her own technical research.
My parents for their support from the beginning.
This research was sponsored by NASA as part of the UA/NASA Space Engineering Research Center. The author gratefully acknowledges the support provided by Dr. Murray Hirschbein and Dr. Gordon Johnston through the grant NAGW-1332. V
TABLE OF CONTENTS
LIST OF ILLUSTRATIONS VI
LIST OF TABLES VIII
LIST OF VARIABLES IX
ABSTRACT 1
Chapter 1: BACKGROUND 2
Chapter 2: INTRODUCTION 12
Chapter 3: DEVELOPMENT OF SPREADSHEET 23
Chapter 4: SPREADSHEET ORGANIZATION 31
Chapter 5: PRIMARY SPREADSHEET EQUATIONS 38
Chapter 6: PROPELLANT DATABASE 67
Chapter 7: SURFACE SUPPORT COMPONENTS .. 75
Chapter 8: IN-SITU RESOURCE UTILIZATION 82
Chapter 9: MODULAR ENGINE COMPONENTS 92
Chapter 10: APPLICATION TO A MARS SAMPLE RETURN MISSION 99
Chapter 11: SUMMARY OF OPTIMUM MISSION PLAN 104
Chapter 12: PROJECT SUMMARY 123
Appendix A: FIGURE-OF-MERIT SPREADSHEET 128
Appendix B: SPREADSHEET EQUATIONS 214 VI
LIST OF ILLUSTRATIONS
Figure 1: FoM Concept 13
Figure 2: Historical Flyby Missions 16
Figure 3: Historical Orbiter Missions 16
Figure 4: Historical Lander Missions 17
Figure 5: Historical Sample Return Missions 17
Figure 6: Lander Missions with Linear Fit 19
Figure 7: Lander Missions with Power Fit 19
Figure 8: Sample Return Missions with Linear Fit 21
Figure 9: Sample Return Missions with Power Fit 21
Figure 10: MSR Mission Date Schematic 26
Figure 11: Refrigeration Unit Heat Transfer 41
Figure 12: Pressure Feed System Schematic 50
Figure 13: Pump Feed System Schematic 51
Figure 14: Feed System Design Selection 53
Figure 15: Pressure Feed Rocket Mass Breakdown 54
Figure 16: Pump Feed Rocket Mass Breakdown — 54
Figure 17: Propulsion Unit Weight Comparison 57
Figure 18: Turbopump Feed System Coefficient 59
Figure 19: Nozzle Geometry 61
Figure 20: Oxygen Plant Test Bed 85
Figure 21: Oxygen Plant Flight Hardware 86 VII
Figure 22: Plant Mass Breakdown 90
Figure 23: Modular Engine Staging 93
Figure 24: Advanced Modular Engine Staging 98
Figure 25: MSR Mission Trajectory 103
Figure 26: MSR Staging Schematic 105
Figure 27: Initial Staging Masses 107
Figure 28: Payload Constraints 108
Figure 29: Mission Variations 109
Figure 30: Mission Mass Breakdown 110
Figure 31: FoM Correlation 112
Figure 32: Sample Mass Variations 115
Figure 33: Return Payload Variations 116
Figure 34: ISRU Comparison 117
Figure 35: Plant Mass Variations 119
Figure 36: Fuel Ratio Effects 120
Figure 37: Nozzle Area Ratio Effects 121
Figure 38: Technology Benefits 122
Figure 39: Technology Mass Savings 123 VIII
LIST OF TABLES
Table 1 - Possible Figure-of-Merit Definitions 15
Table 2 - Distance Factors 18
Table 3 - FoM Program Sheet Outline 31
Table 4 - Initial Mass Component Breakdown 39
Table 5 - Structural Coefficient Dependencies 56
Table 6 - Fixed coefficients 56
Table 7 - Variable Coefficients 57
Table 8 - Pump Feed System Coefficients 58
Table 9 - Pressure Feed System Coefficients 59
Table 10 - Database Propellant Combinations 68
Table 11 - Stoichiometric Combustion Reactions 70
Table 12 - System Chamber Pressures 71
Table 13 - Nozzle Expansion Area Ratios 71
Table 14 - Oxidizer to Fuel Mass Ratio Multipliers 72
Table 15 - Propellant Data Record Fields 74
Table 16 - Modular Engine Masses 97
Table 17 - R-factor Comparison 114
Table 18 - Mission Variation Summary 123 IX
LIST OF VARIABLES
VARIABLE UNITS IDENTIFICATION
Ac Aeroshell Coefficient m2 Area of Nozzle Exit Plane Ap m2 Projected Area A; m2 Surface Area A, m2 Area of Nozzle Throat A1 Cubic Equation Coefficient A2 Cubic Equation Coefficient A3 Cubic Equation Coefficient • c m/s Characteristic Velocity COEFtl Total sum of Coefficients D Determinant for Cubic Equation CIT m Tank Diameter EC Engine Coefficient F kg Rocket Stage Design Thrust FA Fraction of Tank Area Exposed to Solar Radiation FAF Fraction of Fuel Tank Exposed to Solar Radiation
FAQX Fraction of Oxidizer Exposed to Solar Radiation FSC Feed System Coefficient FTC Fuel Tank Coefficient FTSF Fuel Tank Design Safety Factor Gc Guidance Coefficient 2 Gs W/m Solar Flux in Low Earth Orbit g m/s2 Gravitational Acceleration h W/m2-K Convection Coefficient I>P sec Specific Impulse Asp(vac) sec Vacuum Specific Impulse k W/m-K Thermal Conductivity MA kg Aeroshell Mass Me kg Engine Mass Mf kg Fuel Mass
^FS kg Feed System Mass Mpr kg Fuel Tank Mass MG kg Guidance Mass ML kg Payload Mass Mn kg Nozzle Mass MOX kg Oxidizer Mass
M0XT kg Oxidizer Tank Mass MP kg Propellant mass MRATIO Mass Ratio MRF kg Fuel Refrigeration Unit Mass X Mrox kg Oxidizer Refrigeration Unit Mass MRP kg/W Refrigeration Unit Specific Power MRPF kg/W Fuel Refrigeration Unit Specific Power Mrpox kg/W Oxidizer Refrigeration Unit Specific Power Mse kg Structural Mass - Engine Mss kg Structural Mass - Secondary kg Tank Mass M.TP kg Turbopump Mass M,0 kg Initial Mass M1 kg Final Mass kgl/3 MRPRE Refrigeration Unit Mass Intermediate Coefficient NC Nozzle Coefficient Npp Number of Fuel Tanks Noxt Number of Oxidizer Tanks ^SF Nozzle Safety Factor NT Number of Tanks Nu Nusselt Number OXTc Oxidizer Tank Coefficient OXT,SF Oxygen Tank Design Safety Factor O/F Oxidizer to Fuel Mass Ratio
Pc Psia Chamber Pressure PpT Psia Fuel Tank Internal Pressure ^OXT Psia Oxidizer Tank Internal Pressure Psia Pressure Maintained Inside Propellant Tank PF % Fuel Percentage by mass Pr Prandtl Number QCOOL m/kg2* Cooling rate for constant tank temperature Qin m/kg2/3 Heat input to tanks Qsurf W Surface heat flux Qout m/kg2* Heat output from tanks Qi Cubic Equation Intermediate Coefficient RFT Fuel Tank Reflectivity Roxt Oxidizer Tank Reflectivity R1 Cubic Equation Intermediate Coefficient Ra Rayleigh Number FT m Radius of Fuel Tank OXT m Radius of Oxygen Tank T m Tank Radius m Radius of Nozzle Exit Plane m Radius of Nozzle Throat aN m2 Nozzle Surface Area SEC Structural Coefficient - Engine SF Safety Factor SSC Structural Coefficient - Surface Structure S, Cubic Equation Intermediate Coefficient XI
T K Absolute Temperature TF K Fuel Temperature TFAA Thrust Factor Tox K Oxidizer Temperature TSF Tank Safety Factor TI Cubic Equation Intermediate Coefficient 'burn sec Burn Time V m3 Volume VN m3 Nozzle Volume VT m3 Tank Volume VI Intermediate Coefficient V2 Intermediate Coefficient V m3/kg Specific Volume 3 VF m /kg Fuel Specific Volume vox m3/kg Oxidizer Specific Volume WN m Average Width of Nozzle Material wT m Tank Wall Thickness
Greek Symbols
A Absorptivity AD m2/s Thermal Diffusivity «N deg Nozzle Expansion Half Angle P 1/K Volumetric Thermal Expansion Coefficient AV m/s Spacecraft Velocity Change € Emissivity EF Emissivity of Fuel EOX Emissivity of Oxidizer VR Refrigeration Unit Efficiency WRF Fuel Refrigerator Unit Efficiency VROX Oxidizer Refrigerator Unit Efficiency V m2/s Kinematic Viscosity p kg/m3 Propellant Density PF kg/m3 Density of Fuel 3 PFT kg/m Density of Fuel Tank Material 3 PN kg/m Density of Nozzle Material Pox kg/m3 Density of Oxidizer POXT kg/m3 Density of Oxygen Tank Material PT kg/m3 Density of Tank Material A W/(m*-K4) Stefan-Boltzmann Constant CTPR N/m2 Design Stress of Fuel Tank Material aOXT N/m2 Design Stress of Oxygen Tank Material N/m2 Design Stress of Tank Material 1
ABSTRACT
The concept of a Figure-of-Merit (FoM) is developed to assess specific mission designs. The variables for a mission plan are so numerous and interdependent that a single parameter cannot accurately represent the overall design performance. The introduction of in-situ resource utilization (ISRU) and the use of advanced modular engines further complicate the problem. For these reasons, the FoM approach has been proposed to provide a more comprehensive look at the overall picture. The analysis encompasses the important design parameters in addition to the less tangible aspects such as long-term effects, reliability and reparability of the hardware, and the risks that are inevitably associated with new technologies. FoM's have been examined in detail for historical missions and for a proposed
Mars Sample Return (MSR) mission. Results are presented for a conventional MSR mission along with missions incorporating ISRU and modular engines for comparison. It is concluded that this quantitative FoM approach may well become a key tool in the analysis and design of future space missions. 2
CHAPTER 1
BACKGROUND
Planetary missions can be classified into two distinct categories. By far, the most common mission type is represented by a small scale scientific probe, such as either of the two highly successful Voyager spacecraft which journeyed to the outer solar system and beyond.
The probes themselves are designed for extraterrestrial exploration and usually include an extensive array of instrumentation to attain their goals. The common ambition for all of these probes is the further advancement of scientific knowledge. The term "small scale" is used only to differentiate these missions from the second category, which falls under the general heading of "national programs". The Apollo program is a prime example of this second category. A national program is characterized by its scope and complexity. A given program may include multiple missions and require large investments of capital and time. The initial mission architecture for a national program is very extensive, but planning becomes easier and more routine with subsequent missions. Planning for small scale scientific missions is less complicated, but only minor refinements in the overall plan can be made once the probe is launched. Each type of mission presents its own difficulties in the planning stages.
Mission details will be altered frequently in the initial design stages. This should be both expected and encouraged in order to list and analyze a comprehensive set of options.
To further complicate matters, mission details are also subject to various outside influences.
Political, economical, and technological changes may have a profound effect on the planning of a given mission, but are frequently not adequately taken into account in the preliminary planning stages. Instead, these effects are usually realized after much research and planning has already been completed. Significant portions of the project may then result in a loss of
both time and money. Potential changes should be considered and tested in the planning stages so that the project can be designed to handle variations which evolve with the mission strategy.
Political trends will have the greatest impact on national programs and thus directly influence the mission goals. The scientific missions tend to be more isolated from the
political arena, but immunity from all changes is never guaranteed. The political factor having the largest impact on a space program is a national commitment to achieve the specific goals of that program. A national program which has political commitment and public support
usually wins economic support as well. International cooperation may also have a significant impact on the success of a mission. The technological or financial burden may be alleviated in certain specialized areas depending on what resources the collaborator makes available.
If a launch vehicle or scientific instrument is provided by another country, more resources and manpower could be devoted to other areas of the project. Lastly, even political campaigning may have effects on a given space program. A politician who favors a certain mission provides publicity and hopefully garners support for the program. If the politician is elected, political support is guaranteed, and strong public support is indicated also.
Economic trends are closely tied to the political climate and may have immediate or long term effects on both types of space missions. These effects may just modify certain design aspects of a mission, or they could influence all phases and features of mission planning.
Technological innovations are the third outside influence that may affect space missions. Although technological advances are frequently developed for a specific application 4
in space exploration, breakthroughs in other related fields often have a significant impact as
well. For example, an improvement in the technology for producing electrical energy would
greatly influence the space industry. If energy could be produced more cost effectively or
with an improved power to mass ratio, the advances would be utilized and exploited
throughout all related industries. Specific technological progress will be most obvious through
its utilization on individual spacecraft components, but this will in turn require modification
of the mission planning. Therefore, the mission strategy may be changed indirectly through
the use of new technology. More subtle changes may also occur through the application of
already proven technologies. If an established technology is used, but in a newer application, significant design improvements may result.
Any useful analysis tool must be able to easily accept mission modifications to keep
abreast of concurrent changes in the political, economical, and technological areas of the real world. In any case, the changes must be identified and then analyzed to determine their impact in a given scenario.
From the engineer's perspective, technological trends are easier to identify than forecasting political or economic shifts. With this in mind, the technological innovations have
been highlighted and incorporated into the analysis of this research project. Flexibility in mission details has been stressed in order to accommodate any effects due to political or economical changes. The two technological innovations which may eventually provide major improvements in mission capabilities are in-situ resource utilization (ISRU)1,2 and the use of modular engines3,4.
The main premise behind in-situ resource utilization is to use local resources wherever possible to help fulfill the mission needs. This idea is not new since exploration endeavors 5
have historically used local resources to succeed. An analogy to illustrate the historical use
of 1SRU is the discovery of America. The ships under the command of Christopher
Columbus had their supplies replenished upon each landing wherever possible. It was wise
to invest some time and energy gathering the necessary supplies at all of their stops, and
especially at their destination, for the return voyage. If all of the necessary supplies had to
be carried from the outset of the journey, the ships would have been immense. It is
questionable if such ships could have been built, much less if the voyage would have been
successful. Similarly, it would be advantageous to utilize the natural resources available at the
destination of a space mission. Mission planning has thus far been unduly burdened by the
restriction that everything used in the mission must be initially carried from the Earth.
Considering the Earth's deep gravity well for space operations, the concept of ISRU may
provide significant benefits in mission planning.
A natural progression for planetary exploration is to use the same methods that have
been successful in the exploration of our own planet. Historically, the resources most fully
utilized have typically been consumables such as food and water. In order to sustain a
manned presence in the region as lime progressed, other resources such as building materials gained more importance. In more recent times, natural energy sources have become an
important resource due to our society's dependance on electricity. As far as exploration in space is concerned, a power source becomes critical in order to maintain an atmosphere in which humans can survive and be productive.
Because of the advances in transportation technology, exploration has become very efficient in terms of its overall scope and capabilities. The distances covered by today's
planetary probes and the information they send back would stagger the imagination of the 6
previous explorers of our own planet. The only disadvantage is the burdensome reliance on
a propellant source for propulsion. Fortunately, fuels and oxidizers may be obtained from
some planets, moons, and asteroids.5,6 Since such a large percentage of the initial mass of
a conventional space mission consists of propellant7, the utilization of almost any type of
local propellant available at the exploration site could be highly advantageous in terms of
reduced initial launch mass. Although the processed propellant may not always yield optimum
performance, the benefits of ISRU may outweigh any disadvantages in terms of overall
mission planning. Accurate comparison studies must be carried out to determine which
strategies will be most influenced by ISRU.
The argument in favor of ISRU is compelling, but the actual implementation of this
strategy must prove to be both practical and reliable. The main restriction on the widespread
use of ISRU lies in the processing of the extraterrestrial resources. The resources must first
be collected and then either separated from or combined with other materials to become a
useful propellant. In addition, the propellant must be stored in a useful state until it is
needed. In order to maintain this useful state, the energy needs may become a large factor
just for storage alone. The complexity of the production plant depends on the available
resources, their initial state, and the type of processing required to obtain a useful product.
The complexity also depends on the propellant choice as the end product, and the final state of the end product. It is of utmost importance to obtain a balance between the simplicity of
the reactions and the conditions necessary for the reactions to occur. A simple reaction which requires high temperature and pressure to proceed may be less favorable than a more complex reaction which occurs at moderate conditions. The driving design factors will be the total system mass and the reliability of the components. The hardware should be kept as 7 simple and light-weight as possible. A large percentage of the total production plant mass will generally be taken up by the power source.8 Consequently, the plant will become more efficient by reducing its power requirements, and if possible, by reducing the specific weight of the power source. If ISRU hardware can be proven, the concept provides a compelling alternative for increasing the merit of a given mission.
The inclusion of ISRU in mission architectures will provide many benefits for space exploration. The primary advantage would be the preferential reduction of the total initial mass of the space vehicle. In general, reductions would be realized for cases in which the total plant mass is less than the mass of the propellant it will produce. If a plant weighing
100 kg can produce 200 kg of oxidizer through ISRU, it would be more efficient to produce oxidizer at the exploration site than to launch and carry the necessary oxidizer. Because of the exponential decreases in mass involved in orbital mechanics, ISRU may result in a reduction of thousands of kilograms of the initial spacecraft mass for a typical mission to
Mars. Additional benefits could also be realized for longer production times to provide assistance for subsequent missions. For example, a plant which has a projected total production of 400 kg of oxidizer before failure could provide oxidizer for a secondary mission.
The secondary mission could then be launched with an even smaller initial mass. Ideally, if the technology proves reliable, a complete fueling network could be established to further exploration in a very economical and efficient manner.
ISRU technology would provide other benefits by helping to stimulate new research in many related fields. Improved ideas and advances in technology would be forthcoming in fields such as chemical engineering, materials research, and electronic processing control.
These areas of research focus on the chemical reactions, the necessary hardware for the 8
processing, and the remote or even autonomous control of all the process reactions. The
space program continually provides technological improvements for other Fields both directly
and indirectly, and ISRU would be no exception. The development of an oxygen production
plant is a prime example of a technology that will have direct applications on our own planet.
The ISRU concept could provide terrestrial applications for the reprocessing of factory
emissions in the current industrial setting by converting waste carbon dioxide into molecular
oxygen. Currently, the Space Engineering Research Center for the Utilization of Local
Planetary Resources is investigating a nearly identical process in which a plant is being
designed and manufactured to process the predominantly carbon dioxide atmosphere of Mars
into a useful oxidizer.9
Lastly, development of ISRU technology may result in a related increase in economic
growth. History exhibits a trend of economic growth when new resources are discovered and
then utilized. The numerous advantages which may be gained through ISRU warrants a
thorough investigation of the feasibility of the technology. The underlying technology already
exists; it now requires basic modifications and full development of the hardware for use in
extraterrestrial environments.
In order to fully understand the concept of modular engines it is helpful to have a working knowledge of the background and current status of rocket propulsion technology.
Chemical propulsion provides the foundation for all space programs in the world today.
Many other propulsion designs exist, but presently, all working rockets are chemically based.10 The broad heading of chemical propulsion can be subdivided into separate categories by classifying the state of the propellants prior to combustion. They can be a liquid, a solid, or a hybrid of the two. There are advantages and disadvantages to all three, 9
and the final design choice will be dictated by several factors. Chemical propulsion
technology has matured mainly due to a voluminous database established through extensive
testing and numerous launch attempts. Further data is continually being added and updated
from successful as well as unsuccessful launches.
A vast range of other propulsion methods including nuclear, electric, and solar energy
are classified under the general heading of advanced propulsion. Many derivatives based on
these concepts appear to be very promising.11'12 The primary deterrent for the application
of advanced propulsion is the prominent lack of comprehensive test data. Some of the
technology has been adequately tested, but only within limited confines. For example, in the
case of the NERVA project, a specific nuclear engine design was extensively tested, but the
testing was incomplete since it consisted only of ground tests.13 Many of the newer
concepts have not been tested at all, let alone built. The conspicuous fact remains, that to
date, no missions have been successfully launched using something other than chemical
propulsion. There is no doubt that advanced propulsion would clearly provide benefits in
certain areas such as an increase specific impulse. However, sometimes the benefits are
outweighed by an increased complexity and the returns are quickly diminished. Advanced
propulsion could be made to succeed, but the technology demands a substantial commitment
from both the aerospace industry and the government in terms of time and money for
research and development. Unfortunately, the commitment is not immediately foreseeable.
Consequently, the industrial focus is on chemical propulsion and the desired advances
to improve either specific or overall performance of these rockets. Chemical propulsion is
clearly the leading candidate for any planetary missions to be launched within the next decade. Therefore, only chemical propulsion options have been considered in this study. If 10
advanced propulsion is chosen for future scenarios, the mission analysis should still use
chemical propulsion as the baseline for any comparative study to illustrate the driving
motivation and benefits. Chemical propulsion provides the groundwork for any analysis and
therefore provides the foundation for this study also.
The modular engine concept does not just focus on engine design, but also on the
engine integration between successive stages in order to improve the overall performance of
the vehicle. An optimum engine design is one that attains a maximum specific impulse while
maintaining a minimum engine mass. There are several ways to alter the resulting specific
impulse, but variations in the fuel and oxidizer combinations have by far the greatest effect.
Once a specific propellant combination has been selected, other factors such as the fuel to
oxidizer ratio, chamber pressure, and nozzle area ratio which affect the specific impulse need
to be considered. The modular engine concept essentially focuses on simplifying the engineering to one basic design which can supply a whole host of varying thrust requirements.
Larger thrusts are obtained by using a combination of engines in parallel. Engine clusters of various sizes can provide the necessary thrust for each stage of the spacecraft. The benefits
realized through the utilization of modular engines are increased reliability, reduced risks, simpler reparability, and reduced design and production costs. The total mass of the
propulsion system will be somewhat greater for a modular engine design than for one using a conventional design, but the multiple benefits previously stated far outweigh this one liability.
A distinct need has arisen in the aerospace industry for a new method of analysis in
the planning of space missions. Any new analysis methodology must be flexible, powerful, and 11 easy to use. It must be able to include and quantify the advantages of new technologies, and at the same time it must effectively keep up with changing outside influences. 12
CHAPTER 2
INTRODUCTION
A Figure-of-Merit, FoM, is a quantitative number that represents the relative
advantages and disadvantages of a planned mission. The FoM methodology is powerful
because it focuses on overall mission architecture while not losing sight of the importance of
detailed design. The FoM is designed to be flexible because a spacecraft design and its
associated support details undergo continual changes throughout development. The FoM concept readily accommodates any mission changes and identifies which of these modifications would yield optimum results. As the mission becomes better defined, the FoM becomes more
accurate. Unlike other approaches, optimization of specific components will not be individually considered. Instead, the components will be optimized within a larger framework which encompasses all mission details. Because of the high degree of component integration within a spacecraft, the practice of optimizing only one component may degrade the
performance of another. For example, the specific impulse can be increased by raising the chamber pressure for combustion which yields a higher pressure ratio. However, the additional engine mass will reduce the performance returns. The interdependence necessitates the total integration of mission details including orbital mechanics, rocket design, support components, and ISRU components. A graphical representation of the integration scheme is presented in Figure 1. As missions become increasingly complex, there is a greater need to include all possible details to accurately define the mission and identify its merit. The
FoM is most effective when comparing similar missions. It presents an improved methodology for comparing mission strategies by providing sufficient detail to provide significant results. 13 MISSION DETAILS MODULAR - DURATION ENGINES SPACECRAFT DETAILS -AV's Cross-Section /od\ - Structure THERMAL (ooo) - Feed System ENVIRONMENT ' /PI \PQv - Payload
bn AV, STAGING, T,
ROCKET ISRU COMPONFNTS FoM SPRFADSHEFT PERFORMANCE - OXIDIZER PLANT Pc - QUANTITATIVE SYSTEM F - FUEL PLANT SYNTHESIS Ae/A* Ox - FLEXIBILITY __/~N - STORAGE Mf CET "86, PVT. CO., SOLIDS Ms l.T.P, Ml MO \ M, E, R, PRQPELLANT DATABASE Ox = LOX Fuel = CH4 SUPPORT QUANTITATIVE COMPONFNTS SUMMARY PC - POWER FoM 0.87 1.04 0.69 1.16 ISP - THERMAL CONTROL Ae/At
0/F GRAPHICAL SUMMARY Mass Breakdown
M. f.
FoM
Cases
Figure 1: Figure-of-Merit Concept 14
In order for the FoM to be meaningful, it must have a clear definition. A single strict
definition would be useful only for global comparisons. It has been determined that the
definition should remain flexible so it can be tailored to a specific type of mission. Consider
the difference between flyby missions and lander missions. A general comparison could be
made between the two, but if a detailed analysis of different lander schemes is desired, the
FoM definition should be modified to include the individual mass of the lander. For a given
initial mass, an increase in the lander mass would increase the efficiency of the mission and
result in a higher FoM. In the case of flybys, the lander FoM definition would be
meaningless since nothing ever lands on the surface.
It is easy to draw a distinction between technical and economic definitions, and a list of possible examples is given in Table l.14 Technical definitions are more readily definable and verifiable than economic ones. Alternatively, the economic definitions may be more
pertinent to management and planning decisions. The two types of definitions are not as disparate as they may seem at first because there are definite relationships between program cost and total initial mass. Increased masses will invariably result in higher costs. Only
technical definitions have been studied in detail in this project because of the complications involved in gathering accurate cost data. Although cost FoM's have not been included in the
results, the foundation for the calculation of this number has been laid out in the FoM spreadsheet. Cost relationships based on component mass classifications could be incorporated into the spreadsheet to provide an additional FoM for further comparison if desired.
In order to determine the credibility and accuracy of a given FoM definition, historical data has been compiled and analyzed for previous U.S. space missions. The FoM has been 15 calculated for each mission and plotted in Figures 2-5. Mission planning and technology have
-| '"payloadM ^launch
2 Mpa»^ * Distanc9 raveled Mlaunch x Standard distance
g "'sampleM
^LEO
. Total useful mass 4. Launch mass
5 State-of-the-art launch cost of this payload Launch cost of the payload In this mission
6 Total life-cycle cost including development Standard cost
Table 1 - Possible Figure-of-Merit Definitions
improved with time, and consequently, the merit has increased. Initially, it was assumed that the FoM would increase with time and an effort was made to compare all of the missions simultaneously. It was became obvious that large differences in mission planning necessitates the separation of missions into four distinct categories. The four logical categories are flyby, orbiter, lander, and sample return missions. The first three categories used the most basic
FoM definition given in Table 1. The last category relies on definition 3 in the same table which factors in the mass of the sample returned. At first glance, the FoM appears to 16
• Mariner (Venus) • Mariner (Mars) Q Pioneer (Jupiter) • Voyager
1962 63 64 65 66 67 66 69 70 7! 72 73 74 75 76 77 78 79 ^ 95 Mission Launch Date, yeer M pay! oad Note: Figure-of-Merit - M launch Figure 2: FLYBY MISSIONS
200 Lunar Orbiter Missions Apollo Missions Mariner (Mars) Missions 150 Pioneer-Venue Miraion Mars Sample Return i«•> 100 1£ C 60
0 1962 63 64 65 66 67 68 69 70 71 72 73 74 75 76 77 78LfyL 79 N 95 Mission Launch Date, year M payload Note: Figure-ol-Merit - M launch Figure3: ORBITER MISSIONS 17
• Ranger Missions • Surveyor Missions • Apollo Missions • Viking Missions B Mars Sample Return
a
196263 64 65 66 67 68 69 70 71 72 73 74 75 76 77 78 79Q N Mission Launch Date, year Note:Figure-Merit- Figure 4: LANDER MISSIONS
10 • Apollo Miaeione Mara Sample Return
6 -
196263 64 65 66 67 68 69 70 71 72 73 74 75 76 77 78 794 N 95 Mission Launch Date, year Note: Figure-ol-Merit - M6at"^e MLEO Rgure 5: SAMPLE MISSIONS 18
fluctuate with time in all four graphs. Upon closer examination, however, it can be seen that almost all of the decreases in FoM occur because of the increased distances involved in the later missions. Missions which travel to a common body in the solar system, such as the
moon, can be viewed as a separate group within the graph and a relatively linear increase in
FoM over time results. Missions required to traverse greater distances will invariably require more propellant and subsequently have a greater initial mass at launch. To compensate for this, a distance factor (Distance of mission/Distance to Moon) is incorporated into the FoM definition to increase the merit of missions involving greater complexities and longer durations and distances. Definition 2 in Table 1 does just that. With the inclusion of the distance factor, the expected trend of increasing FoM with time became more apparent. Table 2 is a list of the distance factors used in this study.
Destination Distance Factor Moon 1.00 Venus 281.48 Mars 592.87 Jupiter 2024.71
Table 2 - Distance Factors
Curve fits were applied to the different categories of past missions. The primary goal was to predict the FoM for a Mars Sample Return mission by examining the trends of past missions. The correlation constant was derived for both a linear and power fit. Figure 6 shows the lander mission categoiy with the calculated linear equation superimposed. Figure
7 represents the identical data using a power equation for the calculations. The revised FoM definition which includes the distance factor is used, and the resulting trend is plotted on a 19 500- | Surveyor Missions: (R factor - 0.9, D factor -1.0] I Apollo Missions: |R factor - 0.9. D factor -1.0] 4qq 1 • Viking Missions: |R factor - 0.9. D factor - 204] Ti Mara Sample Return (H2-L0X): |R factor 0.4. D factor - 204] 2. I Mara Sample Return (CH4-LOX): [R factor « 0.6. D factor - 204]
300- FoM - -11.055 + 12.391(Launcti Year -1966) i 200- g 100 H duQ 66 70 75 60 85 90 95 2000 Mission Launch Date, year M payload Figure-of-Merit - 0.02536 (R-factor)(10"4)(D-factor) M launch R-factor - Inverse Riek * Reliability' Repairability D-factor - Distance of mission / Distance to moon Figure 6: Lander Missions with Linear Fit
300. I Surveyor Missions: |R factor - 0.9. D factor -1.0] I Apollo MIssions.'IR factor - 0.9. D factor - 1.0] I Viking Missions: ]R factor • 0.9. D factor - 204] I Mare Sample Return (H2-LOX): |R factor - 0.4. D factor - 204] £ I Mars Sample Return (CH4-LOX): [R factor - 0.6. D factor - 204| .t!; 200 _
FoM -16.393 * (Launch Year - 1965)-Q.6565
75 60 65 90 2000 Mission Launch Date, year
M payload 0 5445 Figure-of-Merit (R-factor)(10"4)(D-factor)J- M launch R-factor - Inverse Riek * Reliability * Repairability D-factor - Distance of mission I Distance to moon Figure 7: Lander Missions with Power Fit 20
normalized time line in which the relative year is the base year minus the launch year. The
base year is the year in which the first successful mission was launched in each mission category. In the figure, the FoM value has been placed at the launch year of the specific
mission. Once the fit constants were calculated for the highest possible correlation, the function was extrapolated to span the estimated time frame for a MSR mission. The FoM
for a proposed MSR mission have been separately calculated in the spreadsheet and superimposed for comparison. The mission was planned utilizing two different propellant combinations and both are shown in the figure to be consistent with the historical data. The
MSR details for these calculations will be explained in a later chapter. Figures 8 and 9 are
plots of the sample mission category for both the linear and power fit methods respectively.
The power fit is viewed to be more realistic in all categories since it levels off with time. It is assumed that technological advances will not increase linearly, but will approach a certain limit over time.
Despite the promising trends shown on the FoM plots, the application of this definition has its limitations. The distance factor does not provide an accurate interpretation when applied to voyages to the outer solar system. The distance factor also fails completely when applied to missions that do not have a distinct destination as in the case of flyby missions. An attempt was made to alleviate the shortcomings of the distance factor by using a duration factor instead. The duration factor had identical problems and further complicated the calculations because of the need for a well defined termination date of the mission.
The Figure-of-Merit approach provides many advantages that other analysis methods may not provide. Foremost, the FoM yields quantitative results. Planning decisions for the space program must be based on hard data. Program decisions involve large sums of money 21
• Apollo Missions: |R factor - 0.9. D factor -1.0] • Mare Sample Return: [R factor - 0.4& 0.6, D factor - 204)
40 _ FoM -0.37516 + 2.1011(Launch Year-1969)
SI 20-
66 70 75 60 65 90 95 2000 Mission Launch Date, year M payload Figure-of-Ment - 20.476 (R-factor)(10"4)(D-factor) M launch R-factor - Inverse Risk * Reliability * Separability D-factor - Distance of mission / Distance to moon
Figure 8: Sample Missions with Linear Fit
60 I Apollo Missions: |R factor - 0.9. D factor - 1.0] I Mare Sample Return: [R factor - 0.4 & 0.6. D factor - 204) £ 60
40 - z FoM -2.5493* (Launch Year - 1969)-0.91408 a 20 si II
66 75 60 85 90 2000 Mission Launch Date, year M payload , -0.7375 Figure-of-Merlt - {—Mleiunch— R-factor - Inverse Risk * Reliability * Reparabillty D-factor - Distance of mission / Distance to moon Figure 9: Sample Missions with Power Fit 22 and time, and consequently, cannot be made lightly. Second, the FoM method focuses on the more important fully integrated mission design rather than on specific component optimization. Components must be optimized only up to the point where it starts to adversely affect the mission. By viewing this design problem in a global perspective, the tradeoff point will become more obvious. Third, the FoM strategy enables a uniform method of comparing similar missions. If mission studies are to be accurately compared, the same procedure for analysis must be used to yield meaningful results. Next, the FoM itself is an effective screening parameter for a wide variety of missions. Numerous mission modifications can be quickly and easily changed within the spreadsheet format, and the resultant FoM will indicate whether or not the change was advantageous. A FoM spreadsheet has been developed as part of this project to automatically generate simplified color graphics for fast evaluation of the results. By automatically graphing the results, the trends are easily recognized and can be used for more effective screening. Also, the FoM can be either specialized for specific missions or kept general to apply to a larger base of information. Depending on which type of comparison is desired, the appropriate definition will provide accurate results. Last of all, any definition used is modified by a multiplier termed the R-factor. The R-factor represents a combination of factors for reliability, reparability, and inverse risk. These elements may have a considerable effect on mission planning, but are not directly accountable in the general calculations. A relative approach for the R-factor is used, and every effort has been made to include accurate data while acknowledging the subjective nature of the modifiers. The absolute modifier may not be exactly accurate for a given mission, but when applied through a comparison basis, the relative influences are quite acceptable. 23 CHAPTER 3 DEVELOPMENT OF SPREADSHEET A spreadsheet program has been developed to calculate the Figure-of-Merit for any given planetary mission. Flexibility and convenience were the key factors in selecting a spreadsheet approach rather than a programming language. Spreadsheets provide a simple means for data input which can then be factored instantly into a new FoM calculation. Furthermore, unlike some standard programming languages, a spreadsheet does not require a separate software package to generate informative and complete graphs. Graphical analysis is readily available within the spreadsheet for both specific and general mission details. Important graphs are generated automatically for each new derivation, while additional graphs can be plotted with minimum effort. In addition, a spreadsheet does not require familiarity with any programming languages. This greatly enhances its ease of use and makes it more accessible to the general user. Because a spreadsheet can be compartmentalized in its design, a complete understanding of the entire spreadsheet is not required. To determine the effects of a given design modification, the value within the data cell can be changed without the need to understand how the value is used throughout the subsequent calculations. In this manner, the spreadsheet can be used as an effective analysis tool with only a limited understanding of its overall structure. However, a detailed understanding of the calculations and inherent structure within the spreadsheet can provide a great deal of insight when making modifications to the actual equations or original parameters. 24 A major feature of the spreadsheet is its ability to provide a complete and detailed analysis for all aspects of a given mission plan. Approximately ISO separate inputs are combined through strict governing equations for each FoM calculation in order to produce significant detail. The large number of inputs may at first seem overwhelming to the novice user, but most of the inputs remain unchanged for most mission variations. If major modifications are desired, the ability to measure the effects of these changes has been provided. An effort was made to maintain an appropriate balance between an effective use of engineering approximations and detailed design calculations in the spreadsheet equations. The engineering approximations provide accurate simplifications of the actual equations governing relationships between parameters. Approximations are sometimes necessary when using unknown parameters or rapidly changing technology. The spreadsheet program has been developed specifically for Lotus 1-2-3. Lotus was ultimately chosen as the application software because of its large market share and numerous capabilities. Given the popularity of 1-2-3, many users can easily utilize the spreadsheet program used in the FoM approach. Additionally, Release 3.1 of 1-2-3 is one of the few commercial packages on the market which has the ability to organize multiple sheets into a three-dimensional spreadsheet structure. The structure has been fully exploited since separate sheets are used to organize the spreadsheet and to simultaneously isolate the inputs. Related inputs are grouped into separate design modules which are placed within distinct sheets. This approach is remotely analogous to using subroutines in programming. Placing the input data into distinct modules provides a more convenient and efficient method of identifying data for subsequent modifications. This approach greatly simplifies the overall organization of almost any exceedingly large spreadsheet. 25 A great deal of time and effort from many people has been invested in the development of this spreadsheet. The initial work began in the fall of 1989 at the Uo£A/NASA Space Engineering Research Center in Tucson by Tom Kirsch for Release 2.2 of Lotus. A paper was presented based upon the initial results from this first preliminary spreadsheet.15 The work continued on the refinement of the spreadsheet and related equations by the author, with the full spreadsheet being converted to Release 3.0 to fully utilize its three-dimensional capabilities. The current version is written for Release 3.1 of Lotus 1-2-3 and makes appropriate use of its inherent graphical enhancements. The spreadsheet program has been gradually assembled and continually improved to include a wider range of capabilities. The primary objective of the spreadsheet is to provide an effective comparison of detailed mission studies. To achieve this, the basic structure integrates a side by side comparison of up to four different mission configurations. Emphasis has been placed on flexibility in terms of mission definition and spacecraft design. The full mission is first subdivided into discrete "legs" of the voyage. Any leg is easily modified because its origin, destination, duration, Av's, and mid-flight corrections are all set arbitrarily. Within each leg, there is also the option to include up to four staging maneuvers. The mission schematic for the MSR mission discussed in this paper is presented as an example in Figure 10. The option of using aerobraking in the analysis has also been incorporated even though the technology has not been completely verified. The benefits to be gained through the use of aerobraking seem to outweigh any possible disadvantages and the consensus opinion supports its use on the next major planetary mission.16 Of course, the aerobraking process is dependent upon the proximity of an atmosphere to succeed, and should only be considered for legs with a destination of a planetary body. Conjunction Class Mission, 1999 Opposition Earth Departure Mars Arrival December 1998 September 1999 Earth Arrival Mars Departure September 2001 January 2001 Mars Stopover Time: 485 days Total Mission Time: 1025 days Figure 10: MSR Mission Date Schematic to C\ 27 A strong feature of the spreadsheet is its incorporation of detailed coefficient dependencies for the structural elements and the engine components. It is important to realize that changing one design aspect may have significant effects on other components. For example, by increasing the chamber pressure to yield an improved specific impulse, the propellant tank mass, feed system mass, and thrust chamber mass will all consequently increase at different rates. In the spreadsheet, a given structural element is not assumed to be a fixed percentage of the total spacecraft mass. Instead, the relative mass of a component is calculated from the design variables where possible. Therefore, the mass of each component must be accurately calculated from all of the independent parameters which may affect it. These interdependencies need to be considered in the design stages in order to predict the projected mission capabilities as accurately as possible. The spreadsheet has been designed so that the choice of working propellants and their operating parameters are easily modified. Several propellant combinations should be analyzed for each mission to provide a complete series of comparisons. Since a change in the combination choice may have a profound effect on the mission planning, it is important to select the optimum combination. The chosen spreadsheet package also has the capability to serve as an interface to any supplemental databases. An external liquid propellant database has been utilized in this study to incorporate known rocket engine performance data for a wide range of propellant combinations. Because this data can be readily accessed by the spreadsheet, it saves time and minimizes data storage by eliminating the need for duplication of the figures within the spreadsheet. Additional thermodynamic data has been included within the spreadsheet so detailed storage parameters can be calculated for the propellants either on the surface or in transit. The large database simplifies the process of modifying the 28 rocket design and quickly outlines the benefits as the changes are factored into the mission plan. The rocket design basis used in the spreadsheet has been configured only for chemical propulsion using liquid propellants. Solid propellant options and advanced propulsion options may be incorporated into the spreadsheet and is left for future development. In order to include any type of advanced propulsion, such as nuclear or electric, as an option; the spreadsheet would need to undergo a major restructuring. Because the advanced propulsion options are not likely to be utilized in the very near future, they are not considered in this analysis. Technological advancements using ISRU and modular engines do appear to be forthcoming, so these concepts are supported throughout the spreadsheet program. The spreadsheet includes an extensive treatment of the support components because of the significant impact they can have on a mission plan. In terms of an unmanned mission, the primary support components are the refrigeration units coupled to the storage tanks containing liquid propellants. The units are needed to maintain steady temperatures in the tanks which are below the boiling point of the propellant. The electrical power requirements for the refrigeration cycle are highly sensitive to the tank temperatures that must be maintained. In addition, secondary electrical power is also needed for various surface operations and long range communications. When using liquid propellants, the main power consumption is caused by long term use of the refrigeration units when the propellants must be stored on the surface for extended durations. If refrigeration is also necessary in transit through space, the power consumption is even greater. Generally, if refrigeration is required in the space environment, as in the case of liquid hydrogen, refrigeration units that require a continuous power supply will be included for the entire duration of the mission. Again, the 29 support components are handled separately within the spreadsheet to simplify modifications when design changes are needed. As mentioned previously, the use of ISRU is an important technological innovation for space missions. The spreadsheet recognizes this importance by including separate calculations for the ISRU plant and by providing a simple means to modify the plant design parameters. The plant mass and its associated power source are independently taken into account. If an improvement in the plant technology is realized, the benefits will be immediately apparent from the FoM value. In this manner, the estimated benefits gained through extraterrestrial propellant processing are quantified directly within the spreadsheet. Modular engine capabilities are also available as an option within the spreadsheet. The detailed design aspects of the modular engine are summarized in a separate sheet, and sizing information can be readily calculated for different modular schemes. The primary limitations on the use of modular engines in the spreadsheet is the restriction for use only on return legs and the inclusion of only a select set of propellant combinations. The use of modular engines solely for the inbound return flight is realistic since modular engines typically do not provide sufficient thrust for the outbound legs of a mission. The large initial mass for the outbound legs requires the clustering of too many engines and does not provide significant benefits. By restricting modular use for return only, the maximum number of engines required in one cluster is usually less than thirty. Even this number may seem large in comparison to current designs. The number of available propellant combinations is limited because the modular engine design does not include refrigeration support for the storage of the liquid propellants. 30 A refrigeration option could be utilized, but does not conform to the modular concept which emphasizes a simplified design to increase the reliability and reduce the associated risks. Calculating the FoM through the use of a spreadsheet program has been shown to provide many benefits. The most useful aspect is the simple procedure for revisions of any input value. In general, commercial spreadsheets are a powerful software tool because data can easily be imported from a database, factored into complex calculations, and then yield graphical results. The program has already been configured to take full advantage of the most likely technological advances to occur within the next decade; the use of ISRU and modular engine design configurations. By incorporating a high degree of flexibility throughout the spreadsheet structure, it is hoped that the useful life of the spreadsheet will be extended. 31 CHAPTER 4 SPREADSHEET ORGANIZATION The full three-dimensional capabilities of the application software greatly simplifies the organization of the Figure-of-Merit spreadsheet. The entire spreadsheet consists of 13 separate sheets which are described below in Table 3. The first sheet contains only minimal calculations while the last sheet does not contain any calculations at all. The first sheet is a Sheet Description 1 Calculation Summary 2 Mission Details 3 Rocket Performance Data 4 Rocket Propellant Data 5 ISRU Component Data 6 Support Components 7 Spacecraft Details 8 Modular Engine Details 9 Case 1 Calculations 10 Case 2 Calculations 11 Case 3 Calculations 12 Case 4 Calculations 13 Spreadsheet Map and Variable List Table 3 - FoM Program Sheet Outline condensed summary of the overall results calculated by the spreadsheet. It contains the initial masses for each of the four selected mission cases along with a mass breakdown by category for all of the major components. The R-factor is presented along with the calculated FoM's in a side by side comparison for all four selected case strategies. The last sheet provides a simplified map outlining the spreadsheet structure and includes a complete list of variables 32 for the novice user. Because there is a considerable number of variables used throughout the calculations, the list is helpful for even experienced users. The inputs for each case are grouped into sheets by category. All of the mission details are input into the first eight sheets which precede the four large mass calculation sheets. Most of the user inputs have been separated from the case calculation sheets in order to simplify any modifications to the case calculations. This scheme makes it much easier to locate inputs which are changed routinely throughout the analysis. For convenience, the most commonly modified variables have been confined to one sheet entitled Spacecraft Details which includes most of the component coefficients. Changing a single variable in this sheet may have profound effects on the analysis. For example, if the theoretical aerobrake calculations overestimate the actual experimental results, then the aerobrake coefficient must be increased. Increasing the coefficient will produce an increase in the size and mass of the aerobrake structure, thereby increasing its braking capabilities. Ultimately, the aerobrake will occupy a larger mass percentage of the spacecraft initial mass. Even small details, such as the fuel tank density, are included in the spacecraft detail sheet. While these fine details have much smaller influences on the mission plan, they may have more significant effects on the final FoM value. Because of the nonlinear nature of the spreadsheet equations, it is difficult to determine in advance the relative effects a specific component modification can have on the FoM. Certain values should not be modified on a singular basis because of the interdependencies of the spacecraft variables. For example, if the tank density is changed through the selection of a different material, the maximum allowable tensile stress of the material will probably be modified as well. Interdependent variables of this type have been 33 grouped together to illustrate the possible effects which one minor change may have on related variables. For the most common mission variations, the Rocket Performance sheet contains the variables most subject to frequent modifications. Within this sheet, the propellant combinations to be utilized for each case must first be determined. Other related input parameters that must be selected include the chamber pressure, oxidizer to fuel mass ratio, nozzle area ratio, and vacuum specific impulse. The selection of these variables is not an easy task since such a wide range of parameters may be readily chosen from the propellant database. A supplemental graphical display program has been written specifically for the propellant database to address this problem by helping to determine a reasonable set of initial parameters.17 The corresponding Rocket Propellant Data sheet which follows the Rocket Performance sheet will not require modifications unless an additional propellant not listed in the database is utilized. The Rocket Propellant Data sheet supplies the thermodynamic data associated with each propellant in a tabular format. The internal spreadsheet calculations access these tables automatically. For input values which are not specifically designated in the tables, an interpolation routine could be easily included to provide data at any intermediate point. The storage tank pressures must be chosen from a set of acceptable values because they are the independent variables in these tables. Additional information on the tank pressures that are supported in the spreadsheet is found in Chapter 6. Once the mission plan has been clearly defined, no further changes will be required in the Mission Details sheet which manages all aspects of the orbital mechanics. The total number of legs and corresponding Av's for each are specified in this sheet. The legs are defined using an Earth perspective which begins with the initial outbound trajectory. The 34 spreadsheet will use the legs in the reverse order for the successive iterations in each Case Calculation sheet. A separate input has been provided to direct any mid-course corrections that might be necessary. Details for each mission leg, such as the number of stages and the optional use of aerobraking or modular engines, are also assigned within the Mission Details sheet. The ISRU component sheet will not require any major modifications except when technological advances are made in this area. The ISRU Component Data sheet estimates the mass of each component of the production plant based upon the production requirements. If a change is made to the production rate value, all of the plant components will be automatically resized and the new masses will be calculated. Care must be taken to ensure that the propellant production rate is adequate to supply the return voyage. Within the ISRU Component Data sheet, separate sections using identical calculations have been established for each selected mission case. All of the components for each case are currently configured for the production of oxygen from gaseous carbon dioxide on the surface of Mars. If another type of plant or different hardware is utilized, the calculations within this sheet must be revised accordingly. The sole function of the Support Components sheet is to calculate the mass of any additional surface refrigeration units which may be necessary. Depending upon the mission destination, refrigeration units may be required on the surface of the planet in order to maintain adequate storage temperatures for the propellants throughout the duration of the surface visit. The specifications and mass of the refrigeration unit are based on the net effect of free convection and radiation of heat from the individual propellant tanks. Specific heat transfer properties can be modified from within the Spacecraft Details sheet. The Support 35 Components sheet will need revisions only if the equations governing the heat transfer analysis are modified. If refrigeration is not required on the surface of the planet, the Support Components sheet will not affect the final FoM calculations. The calculations in the Modular Engine sheet are self contained and provide a design summary of the base engine used for this engine option. If the modular option is selected, the modular mass and thrust values must be entered into the appropriate cell locations after they are calculated in the first initial iteration pass. The copying procedure is necessary to avoid circular reference problems for successive recalculations in the Case Calculation sheets. This process calculates the necessary size and capability of the base engine so it can be used in appropriate clusters for the various stages. Any of the engine details which need to be revised separately, should be changed only in the Spacecraft Details sheet. The four Case Calculation sheets contain the majority of the calculations performed within the spreadsheet. Each sheet is dedicated to one of four different mission plans which utilize inputs from all of the preceding sheets. The details for each leg and the stages within each leg are listed for verification purposes. The variables used in the intermediate calculations are grouped under descriptive headings to clarify the organization of the calculations. A mass summary for each stage is included at the bottom of the respective calculation section. Additional regions which contain surface calculations and preliminary mission summaries have been offset within the case sheet to complement the analysis. Three separate summaries have been included to provide different perspectives of the mission plan. The first is a summary of the masses for a pre-destination craft which is followed by a summary of the masses for the complete return craft. The last summary combines the two sets of data and outlines the total mission masses by category. Most of this information is 36 ultimately transferred to cell locations in the Calculation Summary sheet. The only routine changes made in the Case Calculation sheets are related to the feed system choice and are listed under the heading of rocket design details. Because the four sheets have the flexibility to use a different feed system design for each stage, the choices can only be effectively modified using this organizational structure. No changes should be made to other parameters in the Case Calculation sheets unless the basic analysis needs a revision at the equation level. The philosophy behind the design of the spreadsheet is to minimize the work input and maximize the results. This goal has been attained by logically grouping the inputs into a minimum number of sheets to simplify the modification process. Nearly half of the possible inputs for each case have been grouped together in one sheet entitled the Spacecraft Details sheet. In order to capitalize on such complete hardware definition, each detail must be readily accessible and easily changeable. The structure of the spreadsheet has been made to adhere to this concept whenever possible. When a specific design change is desired, it should not be necessary to modify an entire series of inputs. This would complicate use of the spreadsheet and increase the likelihood of possible errors. The associated variables would have a greater probability of not being modified correctly in the input process. Mistakes can be prevented if all of the associated inputs are modified automatically. The spreadsheet incorporates this preferred methodology through the use of "yes" or "no" options. The yes/no options are available for the design options of aerobraking, transit refrigeration for both outbound and inbound flight, surface refrigeration, ISRU, and modular engine implementation. Choosing the "yes" option for modular engines automatically revises the mass calculations for many related spacecraft components. The need for an individual modification of each component coefficient which 37 is directly affected has been eliminated. Similar reasoning is used for all of the yes/no design options. Through this method of simplified modifications, new planning ideas are more easily tested and fewer errors result in the process. The organization of the FoM spreadsheet provides many desirable features for generating comparisons. The greatest strength lies in the program's ability to handle side by side comparisons of up to four different mission configurations called "cases". Each case structure is not unique in its capabilities because a high degree of commonality exists between all four. All of the cases can accommodate the transit refrigeration needs of both the fuel and oxidizer. All four cases can also utilize ISRU for the oxidizer and provide detailed oxygen plant data. Approximations concerning fuel production can be included within the spreadsheet, but no plant specifics will be calculated. Additionally, all four cases can use modular engine configurations on the return flight if desired. The only significant difference between the cases lies in the capability to calculate the surface refrigeration requirements. The spreadsheet has been designed such that only the first two cases can calculate the refrigeration needs for both the fuel and oxidizer on the surface of the planet. Because the heat transfer calculations are lengthy and complex, this capability was excluded from the last two cases in order to maintain a reasonable spreadsheet size. If the basic design structure needs to be reconfigured, temporary or permanent overrides can be placed in appropriate cells throughout the spreadsheet. A comprehensive understanding of the pertinent equations is essential to succeed with this approach. 38 CHAPTER 5 PRIMARY SPREADSHEET EQUATIONS The spreadsheet represents a large set of interdependent equations. At the heart of these interrelationships is a complex cubic equation which solves for the mass of the fuel burned (MF), given the final mass (Mj) of the rocket stage and the change in velocity (Av) of the vehicle. The derivation of the cubic equation begins with a simple equation for the propellant mass (MP) Mp = M0-Ml (1) where (M0) is the initial mass. The propellant mass is a general term used to represent the total mass of the fuel and the oxidizer (Mox). Mp = Mp+Mox (2) In order to provide a comprehensive and detailed analysis within the spreadsheet, the initial mass of the vehicle has been segregated into distinct categories. The major components and their notations are listed in Table 4. Because the mission employs discrete staging techniques, the payload mass (ML) is defined as the sum of all stages to be used, plus the mass of the payload itself. The payload mass for any one stage may include items such as the sample mass, the rover, or even the propellant factory. For a typical staging maneuver, the mass of the spacecraft will be reduced by the mass of the propellant that has been burned. At the end of the burn, the mass of the spacecraft is further reduced by the mass of the engines and their support structure which are then jettisoned. 39 Component Subscript Aeroshell A Guidance G Surface Structure SS Engine E Engine Structure SE Feed System FS Nozzle N Propellant P Fuel F Oxidizer OX Fuel Tank FT Oxidizer Tank OXT Fuel Refrigeration Unit RF Oxidizer Refrigeration Unit ROX Payload L Table 4 - Initial Mass Component Breakdown A more detailed equation for the final mass can be written as: ++ + Afj = +Mg++ME++MFS+MN+MPJ+^ox r ^RF ^ROX ^ Only the masses of the components pertaining to the spacecraft stage will be included in this expression for each individual stage. The first nine terms can be expressed as a function of the initial mass excluding the payload contribution. For example, the aeroshell mass can be written as: MA = ACX[M0-MJ (4) where Ac is the aeroshell coefficient. Using coefficients dedicated for each component, the • 40 expression becomes: Mi = (A^G^SS^E+SE^FS^N^FT^OXT^-M^M^M^M^ (5) To simplify the equation, the sum of the component coefficients is expressed by just one number (COEF-nJ: Af, = COEF^MO-M^MsM^MMX (fi) In order to remove the final mass from the equation, it is useful to utilize the mass ratio equation: M. -Av (7) eiP ' M, ' V*o A value for this ratio is easily calculated since it is assumed that both the Av and specific impulse (Isp) are known. Substituting the previous expression for the Final mass: M, = COEF^-MJ+MsM^Mmx = (8) It should be noted that the masses of the refrigeration units for the fuel and oxidizer are dependent upon the quantities of each inside their respective tanks. A heat transfer approach is necessary to determine a relationship between the refrigeration unit masses and the propellant mass. The schematic in Figure 11 shows the predominant heat flow paths for the various components that are sensitive to heat input within a representative stage. By monitoring and removing excess heat, the refrigeration unit should be able to maintain a constant temperature below the boiling point of the propellant inside each tank. The expression for the required cooling rate (Qcod) will be the difference between the inflow and outflow heat fluxes and is presented in its most general form by using a variable for the 41 Propellant Tank-^ Refrigeration Unit Support Structure Figure 11: Refrigeration Unit Heat Transfer number of propellant storage tanks (NT): (9) qcx1 •= Both heat fluxes are dependent upon the size of the tanks which is directly related to the propellant volume. Using a radiative analysis, the heat input results solely from the solar flux (Gs), and is directly related to the absorptivity (a) of the tank material and its projected area (AP): The refrigeration unit must be sized to adequately compensate for the maximum solar flux incident upon the spacecraft. If the value for the maximum flux is not used in the analysis, too much propellant may boil off and the required Av may not be attained. If the transmissivity is negligible across the full tank diameter, the absorptivity can be expressed 42 solely as a function of the reflectivity of the tank (RT). In addition, the projected area of the propellant storage tanks is clearly calculated by assuming the tanks are spherical with a radius (rT). The equation for the heat influx becomes: Qjn ~ (l-fljOxG^xTtx (rr) (11) Because one or more storage tanks may be shaded from direct exposure to sunlight, it is prudent to include a geometrical shape factor or fractional area (FA): 2 Q* = FA x(l -Rt) x Gt^xnx(rT) (12) Spherical tanks are used in the analysis because of their favorable combination of a minimum surface area for a maximum volume. It is imperative to minimize the surface area to effectively minimize the structural mass of the tanks and the heat absorbed by the tanks. There is also a certain amount of heat transfer out of the propellant tanks through radiation. Using the Stephan-Boltzmann constant (o), the emissivity (e), and absolute temperature (T) of the propellant, the flux is given by: (13) Qcu, = oxexT*xAM By substituting a more detailed definition of the surface area (\) of a sphere and including a safety factor (SF) to allow for a small margin of error, the expression becomes: 2 14 Qom = SFxaxexT*x [4« x (rr) J ( > 43 The separate heat [luxes are now combined to determine the required rate of cooling: N FA R 4 2 Qcoci = T[ V ~ T)G^xSFxoeT ] x[*(rr) ] (15) The radius of the tank can be expressed as a function of the mass it contains by starting with the volumetric definition of a sphere: V = -re r3 (16) 3 3 Mxv = Afrx-^n(rr) (17) a? The specific volume (v) is expressed in terms of the propellant density (p) by: 3Af Ov)3 = (18) 4jip NT 3M (19) r T ~ 4K pNj The mass of the refrigeration unit is calculated as a combination of the cooling rate, the efficiency of the unit (T7R), and the specific power (MRP): MRpy.Q. Af„ = RP^eool (20) nN-xM, r 4 (21) = "" ~"^[FA(l-flr)G -4xSFxoer lf— * T| R I * J^Tip NTi 44 The only true variable in equation (21) is the mass of the propellant since all of the other parameters will be specified in the design stage. From the analysis, the refrigeration unit mass is found to be a direct function of the propellant mass raised to the two-thirds power. (22) Mr = MR (My where MRpre is an intermediate factor defined as: nNTxMBO ,4 [Fi4(l-^r)GJ_u-4xSFxoe7 l (23) —~ 47t pAL In order for equation 23 to hold true, the properties of the propellant inside each tank must be known and kept constant throughout the mission. It is now possible to proceed with the cubic derivation by setting: (24) MFy = (25) MroX MR,P"m {Moxy Equation 8 can now written as: COEFnxM0*il-COEFn)ML+ MRp„r{MFy^MR^toi{Moxy (26) ~ "ratio* Mq To simplify the analysis, it is conducive to use the oxidizer to fuel mass ratio, O/F, which will be specified for each propellant combination. Another similar term, PF, is employed which 45 represents the relative percentage of fuel based upon the total propellant mass. The definitions and interrelationships of O/F and PF are: M ox OX % 100-PF (27) Mr- F % PF The initial mass can now be expressed in terms of the fuel mass. (28) ^OX ^F ~ M0~M1 M, ; x MF+Mp = Af0 1-- (29) Mn f 100-PF Mr 1 + = (L-MNJ (30) PF MB 1+ 1 PF 100 M„ (31) M0=- PF^~MraJ The initial mass is substituted into equation 26, and the resultant equation is: 100M F MR +MR I COEF„x — + (1-CO£F 77n>)Af,+ i PKfm P"ol\r (32) 100 Af„ = Af^x — — PFil-M^ (1 -COEFn)ML* MR +MR-.J (Mpy prtf PT*oi(fl (33) 100 = W^-COEFn) 46 To simplify the equation, let lOOW^-COEFn) VI = (34) and (35) V2 = 1 -COEF,TL Equation 33 can now be written as: (36) MR +MR... (Mfy = V1XMF- V2XML prtr pre(fox To simplify the exponent, both sides are cubed: MR +MR (-^ (M )2 = [VLXM - V2XM Y (37) Pmere f PreOl\ p F F l 2 3 3 3 2 2 MRDre +MR '-' = V1 xM F-3V1 xM fXV2xML prtf prtoxtff (38) 2 2 3 +3V1XMPXV2 XM l-V2 XMI By collecting terms, the expression is reduced to: 3 2 2 tV1]\M F) - MR +MR +3VJ xV2xML WF) (39) pTtr P'*oxm 2 3 +[3VlxV2 xMl](Mp)-[V2 xMl] = 0 47 3V2xM, 2 (Mf?~ MR +MR ^ (MF) [vr- pr*r pr*oi\ f J VI (40) 3V22xM2 V23XMJ L = 0 VI2 Yl3 This equation can be written in the general form of a cubic equation: 3 2 (Mf.) +.A1(Mf.) +j42(Mp)+A 3 = 0 (41) where the constants are defined as: MR,+MR (42) prtr prt(fox 3V2XML - - VI VI V2xM, (43) A2 = 3 VI V2xM, A3= - (44) VI One solution of this cubic equation can be expressed in general terms as: Mf = S.+ T.- — (45) F 1 1 3 Using standard notation, the values of S and T are: (46) Sj = \JRi +}/D 48 Tj - 3/v^ <47> where R is defined as: 3 B 9AlA2-27A3-2A i (48) = 54 and the discriminant, D, is: D = Ql + R] (49) Lastly, Q is given by: ^ 3A2-A\ (50) ^ = 9 Two other roots can be found mathematically, but they will be imaginary. Since the mass of the fuel must be a positive real number, an imaginary solution would be physically meaningless. A choice between two propellant feed systems for the engine design is available in the analysis which will have a direct effect on many of the coefficients used in the stage calculations. The designer can choose either a pressure feed or a pump feed system. The choice is specific to a given spacecraft stage and must be input separately for each stage within each leg. This parameter is located in the Case Calculation sheets, and is the only value that may be modified on a regular basis within the four sheets. The default configuration uses a pump feed system for all outbound stages and a pressure feed system for all inbound stages. If the user is uncertain about which feed system to use, the spreadsheet 49 will automatically generate a recommendation for each stage based upon the required thrust and the calculated burn time. A schematic of a typical pressure feed system is shown in Figure 12. The primary advantages of a gas-pressure feed system are simplicity and reliability. This system uses a regulated high-pressure gas to pressurize the storage tanks and maintain a steady and continuous flow of propellants into the thrust chamber for combustion. Inert gases such as helium and nitrogen are commonly used in conjunction with a bladder mechanism to prevent any mixing between the pressurizing gas and the propellants. Because the storage tanks must be pressurized to levels above the operating conditions in the combustion chamber, the storage tank masses may become prohibitively large for high chamber pressures. As a general rule, the pressure feed system is more efficient for low thrust requirements. The alternative pump mechanism for a liquid propellant rocket engine is the pump feed system. The schematic in Figure 13 provides an overview of the components used in this type of system. The system uses a mechanical pump to draw the propellants out of the storage tanks rather than forcing the liquids out as in a pressure feed system. A centrifugal turbopump is often chosen because of its efficiency in terms of low mass and high pumping capabilities. The pump is usually driven by a turbine which is powered by a gas generator. A variety of schemes exist to supply the gas generator with its working propellants and then vent the exhaust from the combustion chamber. Figure 13 shows a gas generator designed to bleed off a fraction of the rocket fuel and oxidizer as they travel to the rocket thrust chamber. While this process is self-sustaining, it requires a preliminary pressure contribution to initiate the flow. A small high-pressure bottle is included for this purpose. The gas generator exhaust is used to augment the thrust of the rocket by exhausting directly into the 50 Tank Vent Valve Check Valve Drain Valve Gage Pressure Regulator High Pressure Gas Supply Tank High Pressure Gage Gas Valve Gas Bleed Valve Check Valve Tank Vent Valve Oxidizer Tank Filters Drain Valve Filter Restricting Orifice Propellant Valves Rocket Thrust Chamber Nozzle Figure 12: Pressure Feed System 51 Check Valve High Pressure 0= Gas Supply Tank Check Valve Oxidizer Tank t Valve Valve Fuel Pump Oxidizer Pump Gear Case Gas Generator Turbine Valve Valve Rocket Thrust Chamber Turbine Exhaust Nozzle Figure 13: Turbopump Feed System 52 rocket nozzle. Although the added thrust will be minor, the contribution will be beneficial unless it has a negative impact on the nozzle flow characteristics. The pump feed system has the advantage of lower tank masses because the storage tanks are only moderately pressurized. However, the pump feed system adds additional weight not found in a pressure feed system through the inclusion of a turbopump and its associated hardware. In addition, increased complexity is added to the design which may reduce the reliability. Despite these disadvantages, the turbopump is much more efficient for both high thrust requirements and longer bum times. For both low and high thrust requirements, the feed system choice is obvious. It is considerably more difficult to decide which option to use for moderate thrust requirements. For these cases, the best design combination can be determined by using the FoM spreadsheet for a wide variety of configurations. To enable the program to recommend the best feed system, the rocket engine requirements must be initially approximated. The final determination is based upon a comparison of a plot of burn time versus thrust shown in Figure 14. The graph represents a composite of information taken from various references.18,19 If the engine characteristics plot below the graphical reference line, a pressurized feed system is the best design choice. Above the line, a turbopump feed system should be selected. The graph should not be interpreted as a staunch rule for determining which hardware system to use. It should be used as a design aid to provide general guidelines. A plot of this type is very hardware dependent and will change rapidly with advances in technology. Given a graph of the burn time vs. thrust and realizing the relative advantages and disadvantages of the two feed systems, the designer should be able to make an informed decision about which feed system to use. 53 300 ^ 260 O © CO w" 200 160 a> Turbopump E Pressure Feed 100 Feed 3 CD 60 i i i i i i i i i i i i i i i i 11 i ii i i i ill 1—i i i i i 10 100 1000 10000 100000 Design Thrust, F (kg) Figure 14: Feed System Design Selection Typical stage mass breakdowns are shown for both pressure feed and pump feed rocket engine designs in Figures 15 and 16 respectively. Both pie charts use the same initial vehicle mass for comparison. The relative percentages are approximately equal for the majority of the components, while the turbopump mass nearly equals the difference in storage tank masses between the rocket designs. The percentages will vary for different thrust chamber parameters and for vaiying initial masses of the vehicle. The total system mass can be minimized by selecting a pressure feed system for low thrust requirements with low chamber pressures, and a pump feed system for high thrust requirements with high chamber pressures. The structural coefficients are an integral element of the spreadsheet. A concentrated effort has been made to quantify the interdependencies between related structural OXIDIZER 31.0% GUIDANCE 7,0% ENGINE 18.8% FEED SYSTEM 9.1% STRUCTURE 36.9% AEROSHELL 9.5% PAYLOAD 37.7% Figure 15: Typical Pressure Feed Mass Breakdown OXIDIZER 31.0% GUIDANCE 7.6% ENGINE 18.8% FEED SYSTEM 13.8% sTRucTunAL ^S STRUCTURE 36.9% AEROSHELL 9.6% PAYLOAD 3 7.7% Figure 16: Typical Pump Feed Mass Breakdown 55 components. It cannot be assumed that a given structural element, such as the exhaust nozzle, is a fixed percentage of the total spacecraft mass. Instead, the component mass is calculated using the actual design parameters if they are readily available. The component mass calculations are an improvement over the more traditional fixed percentage mass values because the former approach identifies the relative effects of design changes during the initial planning stages of the mission. This results in the greatest accuracy for predicting the projected mission capabilities. In some instances, the fixed percentage approach was used because not all the design variables were available for every component. Table 5 summarizes all of the major structural coefficients and the variables which affect them. The aeroshell, surface structure, and engine structure masses are calculated as functions of the reduced initial mass, which is defined as the total initial mass of the stage minus the mass of the payload. The mass of each of these three components is calculated as a percentage of the reduced initial mass, which does not vary between stages. The coefficients used throughout the spreadsheet for these components are given in Table 6. The guidance coefficient is slightly more complex because it includes a fixed constant in addition to its dependence upon the reduced initial mass. The guidance mass consists of an on-board computer system for communications which is coupled to a mechanical attitude control system for executing course corrections. The guidance system mass will not vary linearly with the reduced initial mass, but will asymptotically approach a lower limit. The computer and mechanical control system must retain a certain size to remain operative. The equation used to calculate the guidance system mass is located in Table 7. Aside from the reduced initial mass, the engine mass will also depend upon the thrust factor required for the rocket stage. The equation used to calculate the engine mass is also 56 Aeroshell = fn(ML, M0) Guidance = fn(ML, M0) Surface Structure = fn(ML, M0) Engine Structure = fn(ML, M0) Fuel Refrigeration = fn(MF, MRPF, NRF, N^, DF, SFBmax, R^, FAF, CF, TF, SFRf) Oxidizer Refrigeration = fn(Mox, MRPox, NRox, N0XT, Dox, SFBmax, Roxr, e FA0X. ox> Tox. SFR0X) Engine = fn(ML, M0) Feed System = fn(ML, M0, Pc) Nozzle = fn(ML, M0, SFnoz, Dnoz, Wno2, c*. Pc, l8pt or) Fuel Tanks — fn(M^i Mq, SFpyi Dpp Ppy, MF, vF, Op-p) V Oxidizer Tanks = fn(ML, M0, SF0XT, D0XT, PQXT' ^OX' 0X' °OXT) Table 5 - Structural Coefficient Dependencies given in Table 7. The coefficient used in the equation is based on existing rocket engine designs.20 Component Coefficient Aeroshell 0.150 Surface Structure 0.320 Engine Structure 0.026 Table 6 - Fixed coefficients 57 Component Equation Guidance MQ = 0.060 * (M0 - ML) + 0.40 Engine ME = Tfact * 0.0177 * (M0 - ML) Table 7 - Variable Coefficients The feed system coefficient calculated in the spreadsheet is a function of the combustion chamber pressure only. The pressure feed and pump feed systems require separate calculations because of the differences in the design and function of the respective hardware. Figure 17 shows distinct trends for the propulsion unit system mass of the two Pressure y Feed * System / CO CO (0 E • Turbopump Feed o System => Q. 2 a. • / y I 1 1 Chamber Pressure, Pc (psia) Figure 17: Propulsion Unit Weight Comparison 58 different feed systems as a function of the chamber pressure.21 The propulsion unit mass is defined as the combined mass of the feed system, propellant tanks, engine, and nozzle. For a pump feed system: Mjt _ Mjp (51) ME MC where MXP is the mass of the turbopump. This relationship is valid because the engine mass (Me) can be expressed as a direct function of the thrust which is in turn a function of the initial mass of the rocket stage (M0). Combining historical data with analytical calculations, the mass of the isolated pump feed system can be expressed as an exponential function of the chamber pressure as shown in Figure 18.22 The exact equation used in the spreadsheet is given in Table 8. The feed system coefficients are based upon a rocket engine which uses a chamber pressure of 500 psia and has a feed system mass equal to 3.85% of the initial stage mass. The choscn percentage is representative of engine designs used in typical launch vehicles.23 TTC b(Pc) FSC = ae - a = 0.0445 b = 7.59 x 10'4 Table 8 - Pump Feed System Coefficients For a pressure feed system, the system mass is a linear function of the chamber pressure as expressed by the equation given in Table 9. Again, the coefficients are based upon a rocket engine which uses a chamber pressure of 500 psia but has a mass equal to 59 80 / 60 Turbopump mass Engine mass / <%> 40 20 l l i i 0 500 1000 1500 2000 Chamber Pressure, Pc (psia) Figure 18: Turbopump Feed System Coefficient 1.28% of the initial stage mass.24 For both types of feed systems, the system mass will realistically vary according to observed trends for any changes in the operating chamber a = 0.0216 ?Co = 500 psia Table 9 - Pressure Feed System Coefficients pressure. However, the initial mass percentage may vary significantly when different hardware 60 designs are employed. The absolute percentages may need modifications, but the equations given will accurately represent relative changes in the feed system mass. The nozzle coefficient is calculated using specific nozzle design parameters and propellant combination performance values. The nozzle mass derivation begins with: MN = PN*VN (52) where pN and VN are the density and the volume of the nozzle respectively. Assuming a uniform thickness for the nozzle, the volume can be calculated using the surface area (SN) and the average thickness (wN) of the nozzle. Mn = PN*SN (53) To simplify calculations, the nozzle shape can be geometrically approximated as a segment of a cone. It is then possible to express the surface area as a function of the nozzle area ratio which is initially chosen. The variables used in the derivation are shown graphically in Figure 19, in which the subscripts 't' and 'e' are used to represent the throat and exit planes respectively. By including the nozzle expansion half angle (aN), the surface area of the cone is: s» - (55) 61 2 n(r,) (56) SM' - 1 sin a N t 1 (57) smaN 2)- When a propellant combination is selected in conjunction with a set of rocket engine operating parameters, it is possible to predict the resultant performance parameters such as characteristic velocity and specific impulse. The throat area can then be expressed as a (e) w, Of, Throat Exit Plane Figure 19: Nozzle Geometry 62 function of the characteristic velocity (c*), design thrust (F), chamber pressure (Pc), and the specific impulse (Isp). c'xF 4- (58) e sp Substituting equations 57 and 58 into equation 53, the mass of the engine nozzle is calculated from: p xc'xFxw N N (59) N P xI sina A. c spX H \ * / In terms of the reduced initial mass and the nozzle coefficient, the nozzle mass is identically: Mn = Ne(M0-MI) (60) A nozzle safety factor (NSF) is included to provide a reasonable margin of safety and the resulting final equation for the nozzle coefficient is: yy _ NspxpwxwNxc t -1 (61) P xl x sina c v w A */ The thrust factor (Tfact) arises in the equation since it is defined as the design thrust divided by the reduced initial mass. 63 The coefficients for the propellant tanks are functions of the tank design parameters and the properties of the stored propellant. The tank mass is calculated from the density and volume of the tank material: Mt = pT*VT (62) Again, the volume can be approximated by using an average thickness and calculating the surface area of the structure. A spherical design is chosen because it minimizes the surface area for a given volume. By fabricating a tank with the smallest possible surface area, the total mass of the tank will be subsequently minimized. Afr = PJ.xWj.XSJ. (63) If the material properties are known, the minimum required wall thickness can be calculated based upon the pressure (PT) maintained inside the storage tank. The basic structural analysis also requires values for the maximum allowable tensile stress of the tank material (CT), the diameter of the tank (dT), and a safety factor (TSF). _ TSFxPTxdT ^ T 4 o The diameter of a spherical tank can be calculated based upon the volume (V) of propellant it must store: (65) 64 The diameter can now be used to calculated the surface area of the spherical tanks. Substituting these expressions into equation 63, the mass of the tank is defined as a function of the specific volume (v) of the propellant by: 3xTspxpTxPrxM xv (66) 2a The tank calculations are identical for the fuel and the oxidizer. The only difference in the calculations will be in the respective values of the mass and specific volume of the chemical spccies in each tank. In order to present the equations in a convenient coefficient form, the tank mass must be expressed as a function of the reduced initial mass. Combining equations 1 and 2: Mox + Mp - MQ-M] (67) (68) = (69) Mp (7-1) KP-MTATIE) 1 Jf _ AM (70) f— 65 Similarly, for the oxidizer: Mox a . V [—+ll (71) i«. The fuel tank coefficient is then given by: 3xTSF*pTxPrxvf(l FT, (72) 2xo 1 + ffl While the oxidizer tank coefficient is given by: OXJ - xPrx^VXVOx(* ~^rurio) 1 (73) 2x0 1 + - (I) The refrigeration unit masses for the fuel and oxidizer will not depend upon the reduced initial mass. Instead, they are primarily dependent upon the mass of the propellant and the temperature necessary to maintain the propellant in a liquid state. The calculation of the refrigeration unit masses has already been derived in equations (9) through (25), and was shown to depend upon a multitude of heat transfer properties. Although the heat transfer calculations are relatively simple, a large number of dependent variables are needed for the analysis. The specific power and efficiency of the refrigeration unit have been mentioned only briefly, but both play an important role in the final mass calculation. The numerous governing equations combine to form a powerful spreadsheet for analyzing the FoM of space missions, but the greatest strength of the FoM spreadsheet over 66 other methods of analysis is its ability to manage supplementary data. By incorporating a rocket propellant database, detailed data on ISRU hardware, and modular engine integration techniques, the program is more flexible and applicable to a broader range of missions types. By integrating all of the available information, the FoM spreadsheet provides accurate approximations for the relative efficiencies of varying mission plans. CHAPTER 6 PROPELLANT DATABASE The Figure-of-Merit spreadsheet can be used to analyze a large variety of missions because it can easily accommodate a wide assortment of propellant combinations for combustion. The aerospace industry often continues to use a propellant combination which has proved to be reliable in the past. However, this is not always the best choice. Although the mainstay liquid hydrogen and liquid oxygen propellant combination provides an excellent specific impulse, it is not the ideal combination for every conceivable mission. In fact, the refrigeration needs for liquid hydrogen impose a large weight penalty which could be avoided through the use of propellants with higher boiling points. Liquid hydrogen is an excellent rocket fuel for escape from the Earth's deep gravity well, but other fuels may be more efficient for interplanetary transfers. While many other fuel and oxidizer combinations have been tested in terms of rocket performance, the comparisons usually stop there. Comparisons must be made in an overall system approach, such as the FoM, in order to adequately weigh the advantages and disadvantages of each specific combination. Consequently, a propellant database has been included so that a large variety of propellant combinations can be analyzed within the spreadsheet for a given space mission. The spreadsheet program can access an external database of propellant performance data which has been compiled within a commercial software package called Paradox 3.0. The database combines both conventional and unconventional propellants which are likely to be found or processed on the surface of the Earth, Moon, or Mars. Table 10 lists the available propellants grouped according to their respective local systems. 68 Earth System H2 + 02 N2H4 + N204 H2 + H202 NH3 + N204 Moon System CH4 + 02 H2 + Al(10%) + 02 SiH4 + 02 H2 + Al(20%) + Oz Mars System CH4 + 02 CO + 02 NHg + 02 CH4 + H202 CO + H202 NH3 + H202 CH4 + N204 CO + N204 N2H4 + 02 CH3OH + 02 C2H5OH + o2 N2H4 + H202 CHgOH + H2O2 C2H5OH + H2O2 CHgOH + N204 C2H50H + N204 Table 10 - Database Propellant Combinations Unconventional propellants, such as methane and silane, have been included to complement the ISRU capabilities of the spreadsheet. If a resources is available, the possible benefits of using it must be considered in terms of increased mission capabilities after factoring in the processing needs. Table 11 provides a list of the stoichiometric equations and the corresponding oxidizer to fuel mass ratio (O/F) for all the propellant combinations in the database. All of the propellant data has been generated from the Gordon and McBride CET86 computer program (Lewis code) that is widely accepted throughout the aerospace industry.25 The CET86 program uses standard thermodynamic equations and relies on the Newton-Raphson iteration method to determine a correct solution. The assumptions made for calculation of the performance parameters include the following: complete adiabatic combustion, zero velocity in the combustion chamber, isentropic expansion with homogeneous 69 mixing, ideal gas law, zero temperature and velocity lags between condensed and gaseous species, and one-dimensional approximations for the continuity, momentum, and energy equations.26 All of the data was generated by running CET86 on a VAX-8650 mainframe. The design parameters for a rocket engine may vary greatly with different propellant combinations. The database has been constructed to manage a broad range of variation by incorporating a complete set of operating conditions while concentrating most of the data about the peak performance values. The specific impulse and the vacuum specific impulse for a given propellant combination are dependent upon the chamber pressure, the nozzle area ratio, and the oxidizer to fuel mass ratio. The performance values for all of the propellant files in the database have been generated by independently varying these three parameters. Consequently, the number of records within each data file will grow exponentially as the number of permutations increases. A reasonable set of values has been selected for each parameter to minimize the size of the database while still maintaining a comprehensive data set that will fully encompass the requirements of most mission variations. The set of values used for each parameter will not be identical for every propellant combination. Instead, the parameters will vary according to the combustion reaction and the most likely operating range for the spacecraft. By configuring the data sets in this manner, a limited amount of data is efficiently used for the most probable operating parameters. For example, the near-Earth combinations include theoretical performance data using a higher range of chamber pressures. The high pressures are necessary to offset the elevated back pressures encountered at sea level during liftoff. For lunar operations, pressures in this range are unnecessary because of the absence of an atmosphere. The limited benefit of an increased specific impulse at higher pressures would be offset by the greater mass of the 70 EQUATIONS O/F MASS RATIO Earth System 2H2 + 02 -» 2H20 8.000 H2 + H202 — 2H20 17.00 2N2H4 + N204 -» 4H20 + 6N2 1.438 8NH3 + 3N204 -• 12H20 + 7N2 2.029 Moon System CH4 + 202 C02 + 2H20 4.000 SiH4 + 202 -• 2H20 + Si02 2.000 10% Al: 243H2 + 2AI + 12302 -»• 243H20 + Al203 7.289 20% Al: 216H2 + 4AI + 11102 216H20 + 2AI203 6.578 Mars System CH4 + 202 -• C02 + 2H20 4.000 CH4 + 4H202 -h- C02 + 6H20 8.500 CH4 + N204 - C02 + 2H20 + N2 5.750 2CO + 202 -• 2C02 + 3H20 1.143 CO + H2O2 - C02 + H20 1.214 4CO + N204 - C02 + N2 0.821 4NH3 + 302 -» 6H20 + 2N2 1.627 2NH3 + H202 - 6H20 + N2 4.313 N2H4 + 02 -> 2H20 + N2 1.000 N2H4 + 2H202 -> 4H20 + N2 4.250 2CH3OH + 302 - 2C02 + 4H20 1.500 2CH3OH + 6H202 -• 2C02 + 10H20 3.188 4CH3OH + 3N204 -• 4C02 + 8H20 + 3N2 10.422 C2H5OH + 302 -» 2C02 + 3H20 2.087 C2H50H + 6H202 -• 2C02 + 9H20 5.131 2C2H50H + 3N204 - 4C02 + 6H20 + 3N2 5.042 Table 11 - Stoichiometric Combustion Reactions combustion chamber required to maintain the pressures. The range and specific values of the chamber pressures for the Earth, Moon, and Mars systems are presented Table 12. The nozzle area ratio is the only parameter which has an identical set of values for every propellant combination. The array used in this set covers a broad range of values 71 Chamber Pressures (Pc). psia Earth System 50,100, 500, 1000, 2000 Moon System 0.1,10, 50,100 Mars System 1,10, 50,100, 500 Table 12 - System Chamber Pressures commonly used in nozzle designs. The nozzle area ratios listed in Table 13 represent the values used consistently throughout the database. Nozzle Area Ratio All Systems 5, 10, 25, 50, 100, 250 Table 13 - Nozzle Expansion Area Ratios The oxidizer to fuel mass ratio for each propellant combination is determined by stoichiometry of the induced chemical reaction. An extended set of mass ratio values is necessary because it may be advantageous in various situations to carry out the reaction in either a fuel rich or fuel lean environment. By using a standard set of multipliers on the stoichiometric values, a comprehensive series of oxidizer to fuel mass ratios is generated. The standard multipliers used consistently in the database are provided in Table 14. The oxidizer to fuel mass ratios available in the database for each propellant combination can be easily calculated by finding the product of the multipliers and the stoichiometric values given in Table 11. The FoM spreadsheet has been configured to make efficient use of the external database. One goal of the spreadsheet is to quickly provide an accurate comparison of a wide 72 0/F Multipliers All Combinations 0.3, 0.7,1.0,1.5, 2.5, 3.5 Table 14 - Oxidizer to Fuel Mass Ratio Multipliers range of propellant combinations to determine an optimum mission plan. The mission plan is not limited to a single propellant combination for each case variation, but allows the utilization of different propellant combinations for both outbound and inbound portions of the mission. The records in the external database files can be accessed from inside the spreadsheet without using a dedicated software database package. A query command on the database will provide all the data that is necessary to complete the inputs in the rocket performance sheet for the subsequent calculation of a case FoM. The retrieved data is not entered automatically, but must be input manually for both the outbound and inbound legs of the mission case being analyzed. The most convenient method is to just issue the spreadsheet copy command to input the data into the appropriate locations at the beginning of the rocket performance sheet. When the database search criterion is entered into the appropriate cells within the Rocket Performance Data sheet, and the query command is issued, the corresponding data is placed directly into the output range located farther down in the same sheet. The input file must first be specified before commencing with any database command. The data within the files can also be modified from within the spreadsheet, but it is recommended that any changes to the files be verified before the actual modifications are. 73 attempted. Familiarity with the 1-2-3 database commands is helpful for complex searches, but the procedure is relatively straightforward in most situations. The propellant parameter data search can encompass the very specific to the very general. Global searches across propellant combinations requires the use of the aggregate data file. However, if only one combination needs to be examined, the smaller, separate files may be used to execute the search routine more quickly. Table 15 lists the ten propellant data fields that exist within each database record. Each propellant file contains approximately 200-250 records. The search criteria range may contain from one to ten values. At least one parameter must be input for each search. Including additional parameters will serve to limit the scope of the search. If all ten fields are specified, only one data record will correspond. In a typical search strategy, the fuel, oxidizer, and chamber pressure will be input. The output from this type of search will be a set of data values for the full range of mass ratios with the data grouped according to area ratios. Essentially, the more preliminary information that is known about the mission, the more effective the search will be. To facilitate the search process, the output from any search is automatically plotted within the spreadsheet to provide an immediate graphical representation of the available propellant data. The graph default settings can easily be revised to accomodate the type of data retrieval. The default graph configuration plots the vacuum specific impulse as a function of the oxidizer to fuel mass ratio. Different families of nozzle area ratios can be supported as necessary. The graph is configured to allow for rapid identification of the optimum specific impulse and emphasize the inherent trends for the data set. The optimum impulse will not necessarily yield the highest FoM for a mission plan, but it does provide a reasonable starting point for a more detailed analysis. In most cases, it will be necessary to 74 Field Variable Units Fuel F Oxidizer OX Oxidizer/Fuel Mass Ratio O/F Chamber Pressure Pc psia Chamber Temperature Tc K Nozzle Area Ratio Ae/A, Pressure Ratio P^e Characteristic Velocity C ft/sec Specific Impulse op sec Vacuum Specific Impulse 6p(vac) sec Table 15 - Propellant Data Record Fields run only a portion of the originally retrieved combinations once the appropriate data trends are identified. The Rocket Propellant Data sheet is a dedicated sheet that has been incorporated into the FoM spreadsheet to manage the supplementary thermodynamic properties of each chemical compound supported in the database. The data is necessary for calculations of the propellant tank size and mass for each stage of the spacecraft. The storage temperatures, densities, and specific volumes for each compound are dependent upon the pressure maintained inside the propellant storage tank. The values of these properties are accessible in the spreadsheet in a tabular format indexed according to the tank pressure. Currently, a limited choice of tank pressures is available until an accurate interpolation routine can be incorporated into the spreadsheet. 75 CHAPTER 7 SURFACE SUPPORT COMPONENTS Space missions which perform special surface operations usually require the use of additional equipment after landing at the exploration site. For an unmanned mission, surface operations include diverse tasks such as surveying, monitoring of environmental conditions, collection of geologic and atmospheric samples, and the execution of scientific in-situ experiments. Special plans and equipment are often necessary to oversee and support the mission requirements that cannot be handled through normal spacecraft operations on the planetary surface. The payioad equipment will need a power source on the surface in order to perform the designated tasks, collect the related data, and transmit the observed results. The term "support hardware" is defined as any non-integral piece of equipment that supplements the mission payioad. A separate sheet detailing the mass and power calculations associated with the support components has been included within the FoM spreadsheet because the additional mass may have a significant influence on the mission plan. For the MSR mission in this study, secondary power is necessary for rover operations, Earth communications, and liquid propellant storage on the surface of Mars. The most critical support components are the refrigeration units because they consume the largest amount of power. The refrigeration units are connected to the storage tanks in order to maintain a temperature below the boiling point of the propellant inside each tank. The refrigeration unit masses are dependent upon the electrical power requirements for the refrigeration cycle, which is highly dependent upon the tank temperature. Therefore, the 76 refrigeration unit masses will vary greatly when different propellants are selected for the return journey. The Support Components sheet accounts for the refrigeration needs on the planetary surface for either the fuel, the oxidizer, or both compounds simultaneously. In general, no additional power will be required for normal communications from the surface since contact is maintained by the spacecraft guidance system throughout the mission. The electrical power needed for communications has already been included in the guidance mass figures. The power unit has already been integrated into the total component mass figures for both the rover and the ISRU production plant as well. As a result, none of this hardware has support components that are handled separately within this sheet. Using this method of accounting, only the refrigeration unit calculations are included within the Support Components sheet. In the event that refrigeration is not required on the surface, this sheet will have no effect on the final FoM calculations. The spreadsheet is flexible and may be organized to provide the most effective method of mass accounting for a particular mission plan. For example, a manned mission could be accommodated by modifying this sheet to manage the extra power consumption. Different types of power sources were analyzed for use by the support components. A nuclear power source was chosen because it provides the most beneficial mass-to-power ratio as compared to radioisotope thermoelectric generators (RTG), solar photovoltaic, and solar dynamic arrays.27 For these options, the technologies are constantly improving and different design approaches should be compared within the FoM spreadsheet before making a final selection for the power source. The electrical and thermal calculations are based upon the specific power estimates for an SP-100 type nuclear reactor design. The SP-100 will meet 77 all of the power requirements in the current mission plan and represents the most recent technology in space reactor design. The specifications and corresponding mass of the refrigeration units are based upon calculations which combine free convection with radiative heat transfer. The analysis employs some approximations, but still provides sufficient detail for realistic calculations. In the calculations, the propellant containment vessel is a spherical shell without any additional insulation layers. The surface storage tanks are envisioned to be flexible structures which use a highly reflective material to provide adequate insulation. The refrigeration units used for transit through space are treated separately and are not considered in this portion of the analysis. Because of the lack of insulation, the refrigeration masses represent a worst case estimate which is practical for initial design stage calculations. The refrigeration requirements can be reduced if a simple insulation scheme is used, and if the transit refrigeration units are used to supplement the surface refrigeration units. The analysis isolates each return stage when calculating the refrigeration unit cooling rate and related mass required for that particular stage. The total refrigeration unit mass that will be required on the surface of the planet is determined by combining all the separate stage masses into one value which is presented in the spreasheet summary. While on the surface, the propellants are stored in the spacecraft tanks which are cooled by supplementary refrigeration units. If both the fuel and the oxidizer require refrigeration, the following calculations are applied separately for each reactant. The diameter of the spherical containment vessels used for each return stage is dependent upon the propellant mass and the tank storage parameters. The number of tanks (NT) within the spacecraft stage is specified as an independent parameter in the Spacecraft Details sheet. The only design 78 constraint is that each set of fuel or oxidizer tanks within a particular stage must be identical. The diameter of the tanks is calculated in the same manner as that used in the Case Calculation sheets for computation of the tank radii. The equation is slightly different because the relation is more useful when expressed in terms of a diameter (d): (74) npxNr where p is the density of the propellant stored inside the tank. The equation used to calculate the heat flux caused by convection is expressed in its most general form as: = h*AAT.-TJ (75) where h is the average heat transfer convection coefficient, and \ is the total surface area. The temperature gradient results from the difference between the temperature at the surface of the body (Ts) and the ambient temperature (T^,). The real challenge in any convection calculation lies in determining the convection coefficient. The convection coefficient can be expressed in terms of the thermal conductivity (k) and the Nusselt number (Nu), which is a dimensionless parameter used to simplify the calculations. (76) 79 For natural convection, the Nusselt number is a function of two other dimensionless parameters; the Prandtl number (Pr) and the Rayleigh number (Ra). The Prandtl number is defined as the kinematic viscosity (v) divided by the thermal diffusivity (ad). The Prandtl number can be calculated for Martian atmospheric conditions as a function of the local temperature and pressure. Since the Martian atmosphere is predominantly carbon dioxide,28 the properties of unmixed carbon dioxide are used throughout the equations as an approximation for the atmosphere. The Rayleigh number is dependent upon the thermal diffusivity and the kinematic viscosity, in addition to the volumetric thermal expansion coefficient (/3), and the temperature difference. For a sphere with diameter (d), the Rayleigh number is calculated by: gxMT,-TJd3 Ra = (77) vxarf An experimental relationship for spheres surrounded by a fluid undergoing free convection has been reported by Churchill.29 .. ^ 0.589x72a4 Nu = 2+ (78) The relationship is valid for: Pr i 0.7 and Fa £ 1011 (79) 80 The convection heat flux can now be calculated from equation 75 because all the values are known. The general expression for a heat flux caused by radiation is: (8#> In this case, Ap is the projected area, a is the Stefan-Boltzmann constant, and e is the emissivity of the surface material. Tj and T2 represent the ambient air temperature (Tamb) and the interior propellant temperature (Tprop) respectively. Defining the projected area of a sphere, the equation simplifies to: 4-t* m <81) As mentioned previously, the net heat transfer is the sum of the convective and radiative heat fluxes. • "r <«) Combining equations 76 and 82: 2 (83) The total refrigeration mass (MR) is calculated using the refrigeration unit specific mass 81 (Mrp) and a given unit efficiency (77): QiufxMgp (g4) MR The mass calculations are based upon the assumption that the units will be kept separate and self-contained within each return stage, rather then using one large unit to manage the refrigeration needs of all of the propellant tanks. 82 CHAPTER 8 IN-SITU RESOURCE UTILIZATION The In-Situ Resource Utilization concept relies upon the local processing of materials or manufacture of goods from indigenous extraterrestrial resources. ISRU is available in the spreadsheet so that its relative effects on a mission can be determined in the preliminary design stages before the mission plan is Finalized. Previous studies of the benefits of ISRU report that significant reductions in the initial launch mass of the spacecraft can be achieved by employing ISRU.30 Because of the mass relationships inherent in orbital mechanics, a small decrease in the payload mass will result in a substantial reduction in the spacecraft initial mass. This savings translates directly into reduced mission costs. The ISRU concept encompasses a wide variety of processes, from the extraction and fabrication of structural materials to the production of oxygen for human life support systems. The most significant benefits are realized when ISRU is used for propellant production. The greatest percentage of any chemical rocket launch mass consists of the propellants used in the propulsion system. An examination of the common propellant combinations reveals that the oxidizer comprises a large fraction of any total propellant mass. Consequently, the spreadsheet includes specific data for the use of an oxygen production plant on the planetary surface. Most of the underlying technology for an oxygen processing plant already exists. A few of the required components are not commercially available, and this hardware must be designed and fabricated specifically for the oxygen plant. All of the components must then be assembled in an innovative and efficient manner. The first step is to build a viable prototype and then demonstrate its reliability. After this is accomplished, the production rate 83 can then be optimized such that the plant mass is minimized while the efficiency of the conversion process is maximized. The specific component masses for the oxygen plant used in the spreadsheet are based on current research at the joint UA/NASA Space Engineering Research Center (SERC) for the Utilization of Local Planetary Resources.31 The Center has been selected to build a full pilot plant demonstration system which extracts molecular oxygen from gaseous carbon dioxide through an electrocatalytic conversion process. The plant is designed to operate on the Martian surface under a variety of conditions. The power requirements and masses used in the spreadsheet are based upon a combination of present and projected plant designs. The minimum production rate that will eventually be demonstrated is 2 kg of oxygen per day. The program uses an analytical scaling model to predict component masses for units requiring greater oxygen production rates. The concept of an in-situ propellant plant that would utilize the Martian atmosphere was envisioned back in the late seventies.32 The feasibility of this innovative approach was later confirmed by the Jet Propulsion Laboratory (JPL).33 The JPL research improved upon the initial overall system design by reducing the total system mass while increasing the system reliability.34,35 In addition, basic thermodynamic and electrochemical models were postulated for the chemical dissociation of carbon dioxide and the subsequent production of oxygen.36 The hardware configuration has matured considerably since the early tests, and the current test bed schematic used by SERC is presented in Figure 20. The cryo-cooler and cryo-vacuum chamber are used in the laboratory to simulate the temperatures and pressures encountered on the Martian surface. The constituents of the Martian atmoshere are presented in Table 9 in decreasing order of their relative percentages. Anaerobic carbon 84 dioxide is currently being used as the supply gas because the atmosphere is predominantly composed of carbon dioxide. A comparable mixture of gases will be supplied to the production plant in later tests to fully simulate the environmental compounds the plant must process. Of course, the equipment that is necessary to simulate Martian conditions in the laboratory will not be necessary on the production unit. The hardware configuration shown in Figure 21 represents a schematic of the actual flight hardware that will be dispatched into space. The included hardware provides the basis for the plant mass data used in the spreadsheet. The laboratory design does not include either the power source necessary for the continuous plant operation or the hardware required for long term storage of the oxygen in a useful state. Both are very important factors in the overall plant design and have been specially integrated in the ISRU Component Data sheet to provide a precise ISRU analysis. Since the spreadsheet considers only the actual flight hardware, the schematic in Figure 21 is used to provide a brief explanation of the current oxygen processing plant design. The atmospheric gas is drawn into the inlet and passes through an electrostatic filter to remove dust particles. The severity of the Martian dust storms will supply the upper limit for the capabilities of this filter. The gases are drawn into the plant network through the use of a compressor. An adsorption compressor would be the most favorable hardware selection, but design modifications are required on commercially available units when using carbon dioxide as the working fluid. After compression, the carbon dioxide is passed through a heat exchanger to recover waste heat from the downstream exhaust flow. An additional heater is then used to further increase the temperature of the flow before the gas is introduced into the zirconia oxygen cell. At these elevated temperatures, the carbon dioxide begins to dissociate into carbon monoxide and atomic oxygen. By passing an electric current through 4 -4— 1.0b 333K 847K cryo 6.4nb cooler 38 68.8kg 68.8kg A 38.9kg ^VWWSAAAr HEAT EXCHANGER CHAMBER 58.8kg 10.0 kg RADIATOR y 1.0b 4- 425K 135GK 68.8kg MEMBRANE 1.0b SEPARATOR XHAUS -Jn— MASS FLOU RATE IN kg>day Figure 20: Oxygen Plant Test Bed Recycle Filter Compressor Compressor Membrane Separator Oxygen Storage Tank Radiator Heater Rgure 21: Mars Oxygen Hant Right Hardware 87 the porous zirconia membrane, the negatively charged oxygen ions are selectively passed through the membrane and recombine to form molecules on the opposite side. For this reaction, the solid electrolyte must be maintained at approximately 1000°C. Other electrolyte materials beside zirconia have been suggested and are being researched and tested.37 This process separation and subsequent collection of essentially pure diatomic oxygen. The oxygen stream is then passed through a radiator for cooling and is then compressed. Depending upon the selection of compressor hardware and capabilities, additional cooling may be required for long term storage. A lightweight and flexible storage container with adequate insulation layers has been used for the preliminary mass calculations within the ISRU Component Data sheet. The waste carbon monoxide and the unreacted carbon dioxide pass through the downstream side of the heat exchanger to heat the incoming gas. Afterwards, the exhaust passes through a radiator and then continues on through a membrane separator. The radiator reduces the temperature of the flow so that the separator will not be damaged. The membrane separator divides the carbon monoxide stream from the carbon dioxide flow. The carbon monoxide is vented back into the atmosphere while the relatively oxygen rich carbon dioxide is recycled and repressurized to the working pressure through a secondary compressor. Recycling is used to recover a portion of the work already accomplished through heating and pressurization, thereby reducing the size of the initial compressor and its component mass. The recycled carbon dioxide flow is combined with the incoming atmosphere downstream of the heat exchanger and the process repeats itself. Oxygen has been produced at rates ranging from 6 ml/min to 27 ml/min using a single cell zirconia test bed at SERC.38 The output gas was positively identified as oxygen by gas chromatography and mass spectrometry. In order to produce 2 kg of oxygen per day, a 88 multiple array of approximately 175 cells will be required assuming the lowest yield cell. The single cell design is currently being expanded and modified to accomodate sixteen zirconia tubes into one unit. Electricity must be supplied to the oxygen processing plant to provide power for the filter, compressors, zirconia cell array, and miscellaneous valves. In addition, the heat exchanger, oxygen adsorption compressor, and zirconia cell array will require thermal power. The proposed power source for electricity and heat is an SP-100 class nuclear generating power unit. The specific power estimates are taken from the latest design specifications of the SP-100 as reported in the most recent literature39. The total output has been scaled down to meet the plant requirements. The ISRU Component Data sheet calculates the plant component masses as a function of the predicted oxygen production rate. If the rate is modified, the component masses and total plant mass will be scaled automatically. The sheet also calculates the required production rate for a successful return flight of the spacecraft from the surface of Mars. The production rale is determined by dividing the oxidizer mass required for combustion in all of return stages of the spacecraft by the available surface production time. For a conservative approach, the production time should only be a portion of the total stay time on the surface to allow for start-ups, shut-downs, or periods of inactivity caused by extended dust storms. The minimum production rate input into the data sheet must be greater than or equal to the calculated rate to ensure that the plant can process sufficient oxygen for the return voyage. A considerable safety margin should be included in these calculations to ensure a high probability for success. 89 The calculations in the ISRU Component Data sheet utilize mass estimates based upon previous work at JPL40 for hardware that is not yet available in the SERC demonstration unit. By combining existing hardware with previous research, the initial mass estimate for the oxygen production plant was 197 kg. The current hardware used in the laboratory is continually evolving through technological improvements in the design and manufacture of the plant components. The power source in particular, is undergoing rapid technological advances. The actual advances in technology since the inception of the project, has already resulted in a reduced total plant mass of 145 kg. The masses of individual components are shown in Figure 22 as a percentage of the total plant mass. It should be noted that the power source comprises a significant portion of the total mass. It is estimated that the total theoretical mass of the processing plant will be reduced to approximately 100 kg by the completion of this project. This value was interpolated from trends observed in the areas of rapid development and represents the mass of an oxygen processing plant with a production rate of 10 kg per day. The projected value incorporates improvements in the oxygen cell and compressor designs only. The cell technology and testing is progressing rapidly and substantial advances have already been realized. The improvements in the compressor design result from the use of an adsorption compressor instead of a typical mechanical compressor. Because, a carbon dioxide adsorption compressor has not yet been adequately tested, only theoretical projections of the improvements are available. Alternate designs have been suggested which do not use an initial compressor in the system as a means of reducing the total mass of the plant. If a cell can operate under lower input pressures, the flow could be accelerated with a fan-type assembly to provide sufficient carbon dioxide flow across the membrane. Care must be taken to ensure adequate flow so that the cell does not 90 become oxygen starved which would impose a degradation in conversion efficiency. Because of the extremely low atmospheric pressures on Mars, more theoretical and experimental work must be performed to ensure that this approach is feasible. Mars Oxygen Production Plant Approximate production rate • 10 kg 02 /day 02 Compressor 15.9% Piping/Storage 6.1% Zirconia Cells 8.9% Membrane Sep. 0.7% Heat Exchanger 2.5% Recycle Comp. 10.4% ,C02 accumulator 6.2% Filter 0.5% Radiators 1.1% Computer 5.2% Mass Margin 2.9% Power Source 39.7% Figure 22: Plant Mass Breakdown The ISRU Component Data sheet provides a separate plant mass calculation for each of the four mission cases in the spreadsheet. This facilitates case comparisons that are based solely upon plant mass variations. A simple multiplier is used in the plant mass calculations within the ISRU Component Data sheet to predict either optimistic or conservative mass estimates. All of the plant calculations are based upon the production of oxygen using Martian carbon dioxide as the feedstock. However, the plant data would still be valid for the 91 case of a lunar mission because carbon dioxide is one of the products from the ilmenite reduction process when using ISRU on the surface of the Moon. Individual plant components would possibly require some minor modifications because of the different operating environment, but the overall design would remain essentially identical. If a completely new plant technology is selected, the entire ISRU Component Data sheet would have to be revised. The spreadsheet structure recognizes the inevitable changes that will occur with time and accommodates them in an efficient manner. 92 CHAPTER 9 MODULAR ENGINE COMPONENTS The ability to use modular engines has been included for analysis in the FoM spreadsheet. The term "modular" refers to physical components which are interchangeable and may be used in conjunction with one another. In the case of rocket engines, it implies that one engine used for a particular stage could be replaced with an identical engine from another stage without any measurable loss in performance. The modular engine concept focuses on both the engine design and the engine integration between successive stages in order to optimize the overall performance of the vehicle. An optimum engine design maximizes the specific impulse while minimizing the engine mass. The modular engine concept simplifies the engineering to one basic design which can supply all of the specified thrust requirements. Larger thrusts are obtained by linking several modular engines in parallel so that the thrust is additive. Engine clusters of decreasing number can provide the necessary thrust for each subsequent, lighter stage of the spacecraft. Figure 23 shows a representative engine schematic which details the engine clusters required for multiple stages in a modular vehicle design. The schematic is based upon the traditional approach of vehicle staging since none of the engines are reused. A spacecraft design which utilizes modular engines may have several advantages over the traditional single engine design. The principal advantage is a reduction in the total mission cost. Compared to past space projects, the overall cost of a proposed mission has become a major factor for gaining public acceptance and governmental support. The expenditures required for the design, development, testing, and production of the engines and 93 Mars Return Payload Cross-Sectional View Modular Engine Leg 6: Stage 1 Interstage 1 En9ine Leg 5: Stage 2 2 Engines Leg 5: Stage 1 4 Engines IAIMA Leg 4: Stage 2 8 Engines Hi s Leg 4: Stage 1 IMPTTTTT 19 Engines Figure 23: Modular Engine Staging 94 spacecraft can be substantially reduced using a modular strategy because it uses only a single engine design. The design, development, and testing processes could be accomplished by fewer personnel over a shorter period of time. In addition, a large production run of many modular engines would be substantially less expensive than smaller production quantities of several unique designs. There are inherent limitations in using just one design to satisfy a large range of requirements, but the tremendous cost reductions will offset the slight penalties in performance. The use of modular engines will also simplify the logistics for assembly and launch preparation. Their small size will allow for routine transport of the engines from the production facility to the assembly location. Historically, as the engines become more massive, the transport problems compound. Furthermore, if engine damage is detected upon final assembly, the problematic modular engine could easily be replaced from standard inventory. Dedicated spares and extra hardware would no longer be required. Even launch preparation would be simplified because training would be limited to one engine type only. The preparation of each engine would be identical to the preparation of all others, thereby lessening the possibility of mistakes caused by human error. Other benefits which could be gained through the use of modular engines are indirectly cost related. There is no doubt that a modular engine design would improve both the reliability and reparability over traditional design approaches. With only one engine to test and qualify, the testing process could be expanded and made more rigorous to increase the accuracy and reliability of the resultant data. Furthermore, repairs will quickly become routine by selecting a straightforward engine design. The increased reliability and reparability will be apparent in a larger R-factor for each mission case that utilizes modular engines. The 95 R-factor accounts for even the relatively intangible aspects of a chosen design strategy in an overall system evaluation. The predicted reliability is likely to have a significant impact on the mission evaluation in the preliminary design stages. Unfortunately, the reparability is difficult to evaluate until specific problems actually occur. At this later point, the issue of reparability can assume a level of damaging proportions. The preferred tactic is to consider the reparability in the early planning stages in an effort to predict and plan for realistic solutions to possible operational problems. This is more easily accomplished through the use of modular engines. The modular engine design would provide many advantages over traditional design methods both before and after liftoff. Within a modular design, the failure of a single engine will usually not induce a total mission failure. For example, in a cluster of ten modular engines, the failure of a single engine would still leave ninety percent of the thrust available to complete the objectives. If desired, an additional engine could be included in the modular design to account for a possible failure. The increased mass would be minimal, and the greater reliability for the mission could make the inclusion worthwhile. Alternately, the failure of an engine in a staging design that uses only one large engine will be catastrophic. Including an extra engine in the traditional design case would substantially increase the initial mass and would not be practical. The redundancy available from the modular strategy is inexpensive and easily implemented, and should therefore be included in the early planning stages for all future missions. The ability to use modular engines has been included directly in the spreadsheet. The spreadsheet has been designed so that modular engines can be used only for the inbound return flight. This is a realistic limitation since modular engines typically do not provide 96 enough thrust for the outbound legs of a mission. The thrust capabilities and mass of the modular engines are scaled according to the final thrust requirement upon return to LEO. Fueling engines from the same propellant tanks, would require multiple feed lines and hardware duplication, thereby increasing the complexity of the rocket design. Instead, each engine is integrated as a discrete package which greatly simplifies the design. Each engine will function independently of the others, with only the electronic controls linked. The modular engine scheme will have a higher total mass for the propellant tanks than a conventional design, but this mass increase will be partially offset by the savings in nozzle weight. It should also be noted that the masses are based upon a pressure feed type liquid propulsion design. The pressure feed system provides a simple yet effective design for small thrust requirements. The spreadsheet program calculated and summarizes the design details for the modular engine to be utilized in a given mission. The Modular Engine sheet details the mass of each subcomponent in each of the four mission cases. Within each case, an identical engine design is used, but for different cases, the design and total mass of the engine may vary. An example summary is given in Table 16 which lists the basic engine subcomponents and their masses for a design using a methane and liquid oxygen propellant combination. For a typical MSR mission, the masses are relatively small because of the small payloads which determine the size and capabilities of the engine design. Identical engines will be used in clusters of various sizes for all of the return stages. The total initial mass of the spacecraft can be reduced even further through the use of an advanced modular engine scheme which would reuse engines in successive staging maneuvers whenever possible. The propellant tanks would be jettisoned along with the 97 Subcomponents Mass fka) Fuel Tank 0.25 Fuel 0.59 Oxidizer Tank 0.25 Oxidizer 1.64 Feed System 0.11 Structure 0.11 Thrust Chamber 0.12 Nozzle 0.26 Total 3.33 Table 16 - Modular Engine Masses extraneous engines after each ignition. A cross-section of the spacecraft stages required for a return voyage from Mars is shown in Figure 24. The numbers associated with each engine in the diagram represent the number of times the engine will be fired. The number also indicates the stage number of the structure with which the engines and tanks will be jettisoned. For example, all nineteen engines will be used in the initial launch of the spacecraft shown in Figure 24. After the first bum, the engines and tanks labelled with the number one will be discarded along with the interstage structure. The inner eight engines will all be used again in the next thrust maneuver. The process continues until only the last stage remains. In this scenario, the engines are fired no more than five times each. Care must be taken to position the engines symmetrically about the payload centerline to ensure uniform thrusts with no out of plane components. Further work is necessary to quantitatively estimate the mass savings that could be achieved through an advanced modular staging technique. Currently, only the conventional modular staging approach is supported in the FoM spreadsheet. iffc O ft 98 O © xijy O / Rocket engine lO^OIO Additional Propellant tanks l?fc o©0© o © Stage 1 o ©ofo ©2® o © 0 o Stage 2 Stage 3 ©o© o Stage 4 Stage 5 Rgure 24: Advanced Modular Engine Staging CHAPTER 10 APPLICATION TO A MARS SAMPLE RETURN MISSION An unmanned Mars sample return (MSR) mission has been chosen to illustrate the use of the spreadsheet and reveal insights that are gained through its use. The FoM spreadsheet is a powerful analytical tool, and its strengths are most evident when a comprehensive array of results is examined and compared. A successful Mars mission would provide a multitude of benefits which have been described at the annual Case for Mars conferences over the past several years.41,42 A policy statement released from the Case for Mars IV working group succinctly summarizes the five major areas that would profit from a Mars expedition.43 Scientific and educational fields as well as the more general social, economic, and international areas would all benefit. The possible benefits have been outlined in detail for each area by the working group, and the main reasons are summarized here. A wealth of scientific data could be collected from Mars concerning planetaiy evolution, climate change, and the origin and evolution of life. The exploration of Mars as a national endeavor would set new goals for education as a vehicle for promoting scientific and technical literacy. The social motivations would fulfill man's natural drive to explore and expand while developing new opportunities for human advancement. In economic terms, a Mars program would provide a channel for maintaining the national technical capabilities and establish a constructive focus for future human enterprise. Finally, Mars exploration would open a new avenue for diplomatic interaction between spacefaring nations and provide a natural leadership role for the United States in 100 world affairs. These five benefits provide a compelling argument for the vigorous pursuit of future Mars activity. Although a great deal of information about Mars has been accumulated through remote observations, actual landings are vital to further advance our knowledge. The two Viking lander missions were very successful and provided a wealth of scientific and technical data that is still used for analysis today. While, the Viking missions rapidly advanced our understanding of the planet, they also left many unanswered questions. The possible existence of life on Mars is still debated because of ambiguous results from the surface tests. With our greater understanding of Mars and advances in space technology, future experimental missions could provide answers to more specific questions. Mars exploration has been proposed for many years, and numerous papers have suggested specific plans for both manned and unmanned missions. Many innovative and diverse ideas have been proposed that rely upon both proven and new technologies. Further analysis should build upon this earlier work which defines a baseline for comparison studies. A typical baseline mission has been outlined by combining earlier studies and has been duplicated within the FoM spreadsheet. An effective strategy is to focus on the most promising mission architectures and modify parameters until the FoM is maximized. An increase in the FoM over the baseline mission, indicates clear advancements. Negative results are also useful because they identify detrimental trends that should be avoided in the initial planning stages. The MSR mission has been selected for this analysis because of its current base of scientific and industrial support and its high probability of occurrence. Proposals for more thorough Mars exploration were endorsed by the recommendations of the National 101 Commission on Space which called for the establishment of a bridge between worlds and a more vigorous approach to the settlement of space.44 Additional support was garnered through a supplementary report by Dr. Sally Ride which advocated a more detailed investigation of the Martian surface as one of four points outlining a comprehensive plan for the future of the space program.45 With the Space Exploration Initiative (SEI) set forth by President George Bush on the twentieth anniversary of the historic Apollo landing, the exploration of Mars as a goal of the U.S. space program gained even more momentum.46 Other prominent reports have also called for increased Martian activities.47,48'49 Each of these studies either make a direct recommendation for a MSR mission in their strategic proposals, or would benefit directly from the results of a MSR mission. The most recent report by the Synthesis Group outlines four options to achieve the SEI objectives, including one options which focuses on Space Resource Utilization.50 It is highly recommended that a robotic mission should first be initiated as a means of qualifying critical ISRU hardware before a manned mission is attempted. A MSR mission which uses ISRU would demonstrate the reliability of this new technology as well as provide invaluable scientific data. If a dedicated commitment to a MSR mission is announced, the design and subsequent production of the spacecraft and its payload could begin almost immediately. All of the necessary technology already exists. The primary task to be accomplished before the launch of the mission would be the testing and qualification of the flight hardware. A low energy, conjunction class orbital transfer requires the least amount of propellant, and could be used to minimize the initial launch mass of the vehicle. The higher energy, opposition class mission has not been selected in this study because time is not an important factor in a robotic mission plan. For an unmanned mission, the effects of radiation exposure and zero gravity 102 which accompany interplanetary travel can be ignored. If an ISRU option is chosen, the conjunction transfer would provide adequate stay and production times on the surface of the planet, while the opposition transfer would not. The conjunction class plan is more favorable for a MSR mission and could be completed in less than three years from the time of Earth departure to Earth return. The changes in velocity (Av's) required by the various spacecraft stages for orbital maneuvers essentially represent the sum of the energy expended to complete the mission. The Av's that have been input into the FoM spreadsheet for the mission correspond to a surface stay time of 1.3 years and a total mission duration of 2.8 years. A schematic for this mission including the important dates is shown in Figure 25. It should be noted that by using nominally higher Av's, the surface time could be reduced to 1.0 years and the total mission duration could be reduced to 2.5 years. However, a decrease in the mission duration will always result in an increase of the initial launch mass of the spacecraft. The FoM spreadsheet will enable the determination of the optimum tradeoff point between initial mass and duration on a strictly technical level. A Mars mission has been chosen for illustration, but flexibility is a major strength of the FoM spreadsheet. A simple change in the appropriate Av's will reconfigure the spreadsheet to designate the Moon as the new destination. Under the proposed SEI plan, both Mars and the Moon are the leading candidates for destinations of the U.S. space program in the near future. 103 Conjunction Class Mission 1999 Opposition Earth Arrival September 2001 Mars Earth Sun Earth Departure December 1998 I Mars Arrival September 1999 Mars Departure January 2001 % Mission craft on Martian surface Mars Stopover Time: 485 days Total Mission Time: 1025 days Figure 25: MSR Mission Trajectory 104 CHAPTER 11 SUMMARY OF OPTIMUM MISSION PLAN The optimum mission plan must minimize the initial mass of the spacecraft in LEO and minimize the risk, while maximizing the reliability. Because the program cost is directly proportional to the initial mass, minimizing the initial mass will effectively minimize total expenditures for the project and raise the resultant FoM. The risk, reliability, and reparability are all factored into the FoM calculations. Therefore, the mission plan with the highest calculated FoM will result in the most practical and efficient solution. The staging strategy used for each mission case is identical and is shown in Figure 26. The MSR mission has been subdivided into six different legs, with four of these legs using multiple stages. Both of the second stages of the interplanetary transfer legs include allowances for mid-course corrections. The correction allowances could be safely reduced by a factor of up to 100 because of the increased precision in the prediction of orbital trajectories. A considerable amount of extra fuel is included in the preliminary design stage to provide a liberal safety margin. The mission begins in LEO and uses energy efficient Hohman transfers to maneuver the spacecraft into a circular Mars Parking Orbit (MPO) at a height of 200 km above the mean surface of the planet. Mars capture is accomplished through a combined maneuver of aerobraking and propulsive braking. The lander vehicle and operational hardware later descend to the surface from this circular orbit. After surface operations are complete, the return vehicle ascends to MPO with the planetary sample. Another Hohman transfer is then used to initiate the interplanetary journey back to Earth. To minimize the propellant requirements and the total spacecraft mass for the return trip, the Earth Departure Mars Arrival Hohmann transfer to Mars LEO to Hohmann Parking Orbit transfer (MPO) towards Mars Av = -2.7 km/sec Av = 3.8 km/sec Mid-course correction Av = -0.5 km/sec MSR Vehicle Leg 1 Ascent to MPO Av = 3.5 km/sec Leg 6 MPO to Mars Soft Landing MPO to Av = -1.0 km/sec Hohmann Hohmann transfer transfer to highly eccentric towards Earth Earth orbit Av = 2.7 km/sec Av = -1.0 km/sec Mid-course correction Av = -0.5 km/sec Earth Braking Mars Departure Current Staging Parameter Leg1:2stages Leg4: 2stages Leg 2: 2 stages Leg 5: 2 stages Leg 3: 1 stage Leg 6: 1 stage Figure 26: MSR Staging Schematic 106 craft is injected into a highly eccentric Earth orbit which has a periapsis of 500 km. Once again, both aerobraking and propulsive braking are used for the capture maneuver. Conservative figures for the aerobrake performance have been used because of the uncertainty and lack of true experimental data for this technology. Figure 26 also lists the Av's required for each leg of the mission plan. The net Av required to complete the MSR mission is 15.7 km/sec. This plan assumes that a limited space infrastructure will exist to provide routine transfer operations in Earth orbit. Both the launch and capture of the MSR spacecraft are dependent upon other space vehicles for the mission to succeed. Different plans and options must be analyzed if the infrastructure will not exist by the time of the predicted mission launch date. The initial mass of each subsequent stage in a multistage rocket decreases at an exponential rate because of the inherent relationships in orbital mechanics. Specific mass values for all of the stages of a MSR mission are given in Figure 27. The data represents a mission which uses a methane and oxygen propeilant combination and includes ISRU. An oxygen production plant is placed on the surface of Mars, and produces enough oxidizer to successfully return a 1 kg Martian soil sample to Earth. Therefore, the payload mass for the last stage is 1 kg, and the initial mass of this same stage is shown to be 6.4 kg. The last mission stage is only a portion of the payload launched from the Martian surface. In order to carry this full payload, the required initial mass of the Martian launch vehicle must be 162 kg. The mass effects continue to propagate and increase with each previous stage, and the required initial mass which must be launched from LEO exceeds 38,000 kg. Other propeilant combinations show a nearly identical trend in the successive stage masses, and the figures are comparable. 107 CH4 + 02 Propellant Combination Return Sample • 1 kg 100000 a 38026 y/S'I 8tag« 1 l l Stag* 2 10000= 4886 2 1000= 770 7. to 8 162 2 100 = 7. CO 24 10 = 6.4 7. wi — "m ' '|" 'I' 2? Leg #1 Leg #2 Leg *3 Leg #4 Leg #6 Leg #6 Stage Number per Leg Figure 27: Initial Staging Masses As discussed in Chapter 6, a propellant database can be linked to the FoM spreadsheet to effectively analyze a large variety of possible propellant combinations for a given space mission. By using different propellant combinations, the mission characteristics may vary significantly. For example, a fixed initial mass in LEO will limit the amount of payload that can be returned from the surface of Mars for various missions. Figure 28 summarizes the sample masses that can be returned for varying spacecraft designs using four different propellant combinations constrained by an identical initial mass in LEO. Assuming that the Space Shuttle will be used as the primary launch vehicle for the MSR mission, two separate launches will be required launch a vehicle capable of returning a significant sample 108 Dual Space Shuttle Launch Total Payload Allowance - 46,460 kg 3.0 | ^ H2 * 02 1=1 H2 * 02 (ISRU) BS CH4 • 02 (ISRU) ESS H2«02/CH4*02 (I8RU) 10 kg/day 1.1 kg/day Oxygen Production Rate Figure 28: Payload Constraints mass. Therefore, the initial mass of the MSR spacecraft is fixed at 45,460 kg, which is twice the proven payload capacity of the Space Shuttle. Two distinct data sets are presented based upon the predicted oxygen production rate of the ISRU plant. The plant is currently being designed to produce 10 kg/day, but a much lower rate would optimize the MSR mission. A rate of 1.1 kg/day would be adequate for a mission in this category. The mass reductions resulting from a decreased production rate are also presented in Figure 28. The production rate has no effect on the first propellant case because ISRU is not used in this mission plan. The more energetic propellant combination of hydrogen and oxygen is burdened by the need to carry fuel refrigeration units throughout the mission. Consequently, the methane and 109 oxygen 1SRU case is capable of returning a sample mass which is approximately 20 times as great. Therefore, high performance propellant combinations must also have an efficient overall system design in order to successfully return substantial sample masses. A more effective method of examining the mission results is to specify a fixed value for the sample mass to be returned from Mars. The initial mass in LEO is now the dependent variable, with the mission efficiency being inversely proportional to the initial mass. For a fixed sample mass, the FoM is an excellent indicator of the mission efficiency. The four mission cases presented in Figure 28 have been analyzed for the return of a 1 kg soil sample from Mars. The FoM values for each case are given in Figure 29. To provide further detail, the resulting initial masses are itemized by specific components in Figure 30. The primary Mars Sample Return Maes - 1 Kg Oxygen Plant Production Rate - 10 kg/day 0.35 H2 * 02 H2 * 02 CH4 • 02 H2 * 02/CH4 * 02 (I8RU) (IBRU) (I8RU) Propellant Combination Figure 29: Mission Variations All-ear th-carrled LOX/H2 (Mast at LEO - 69,170 kg) FUEL 6.6% REFR OX 4.1% AE R03HELI REFR FUEL 7.0% ,_ -- ENGINE OX 47.0% ROVER OTHER 6.2% FUEL TANKS OX LEAKED OX TAN KB FUEL LEAKED GUIDANCE 0.0% NOZZLE 6.0% STRUCTURAL 6.7% FEED SYSTEM 6.2% Earth-carried H2/LOX (I8RU) (Mass at LEO - 06,360 kg) FUEL 0.3% OX PLANT AE R08HEL L REFR OX 4 .3 _ OX 61 6% E NQINE ROVER OTHER 0.0% FUEL TAN K6 OX LEAKED OX TANKS FUEL LEAKED GUIDANCE fl.4% NOZZLE 6.3% STRUCTURAL 6 2% FEED SYSTEM 0.0% Earth-carried CH4/LOX (I8RU) (Mast at LEO • 37,690 kg) FUEL 17.0% OX PLANT REFR OX 4.7% AEROSHELL ENGINE OX 47.6% OTHER 60% HOVER FUEL TANKS OX LEAKEO OX TANKS GUIDANCE 6 0% F UEL LEAKED NOZZLE 6.6% STRUCTURAL 6.6% FEED SYSTEM 8.2% H2/LOX OUT WITH CH4/LOX IN (IBRU) (Mass at LEO • 61,930 kg) FUEL 0.6% REFR OX 4.3% OX PLANT ^ REFR FUEL 0O%__^-- AEROfiHELL E N01NE OX 47.2% ROVE R OTHER 6.4% FUEL TANKS OX LEAKED OX TANKS GUI DANCE 6 0% FUEL LEAKED NOZZLE 6.6% STRUCTURAL 6.7% FEED SYSTEM 0.2% Figure 30: Mission Mass Breakdown Ill differences between the mission cases result from the required inclusion of a fuel refrigeration unit when hydrogen is used as the fuel for the spacecraft in cases 1 and 4. The fuel refrigeration units comprise approximately 8% of the initial masses for both cases. In addition, when hydrogen is used as the fuel instead of methane, the amount of fuel leakage during the mission is significantly higher. Hydrogen is an excellent fuel choice in terms of chemical kinetics, but these other problems make it less favorable for long duration missions. It should also be noted that the oxygen production plant masses for the ISRU missions are less than one percent of the total initial mass of the spacecraft. Figure 31 shows a direct comparison of the vacuum specific impulse (I^vac))' the initial mass in LEO (MLEO), and the resultant FoM for 11 different fuel combinations used in identically planned MSR missions. The data is presented in an approximate decreasing order of the vacuum specific impulse. Hydrogen is the most energetic fuel choice and is listed first, while carbon monoxide supplies the lowest performance and is presented last. The graph shows no discernible relationship between the 1^^ and the FoM, but clearly shows the inverse relationship between the FoM and the initial mass of the spacecraft in LEO. Therefore, the selection of specific impulse as the prime indicator of the relative merit of a mission plan is neither accurate nor adequate. The FoM concept alleviates this problem by encompassing more of the important mission details in its definition and subsequently integrates these aspects in its quantitative calculation. By concentrating on the FoM values in Figure 31, the choice of methane as a fuel is seen to provide comparatively large FoM's in relation to the other fuel choices, while using either oxygen or nitrogen tetroxide as the oxidizer. The methanol and oxygen propellant combination is also highly favorable, but when the identical fuel is instead combined with nitrogen tetroxide in the thrust chamber, this 112 Mars sample return mass = 1 kg o 0 to vs "cT 1 q! to 1420 Ulto 400 ££ 13 HI O to mm LL1 3 -J O Oxidizer = 02 1 Oxidizer = N204 '•mm CH30H CH4 NH3 C2H50H Fuel Choice Figure 31: Specific Impulse Comparison 113 propellant combination results in the lowest FoM for any mixture shown in the Figure. This is primarily caused by the exceptionally large O/F mass ratio that is required for a stoichiometric combustion reaction of methanol and nitrogen tetroxide. Because the specific mass of the nitrogen tetroxide is much larger than the specific mass of the methanol, the O/F mass ratio should be kept to a minimum in order to minimize the required initial mass of the combined propellant. In Figure 31 and all subsequent figures, the FoM is shown to have a high degree of sensitivity. A small theoretical increase of just 0.1 in the FoM for the methane and liquid oxygen combination translates into a mass savings of approximately 26,800 kg when analyzing sample masses on this order. This represents a substantial 39% reduction in the initial mass. Therefore, any methods or techniques which produce even slight increases in the FoM should be analyzed and exploited. Numerical values of the FoM for the different propellant combinations are given in Table 17 along with the associated R-factors used in the calculations. The use of ISRU provides numerous benefits in mission planning with the primary advantage being the substantial reduction of the initial spacecraft launch mass. In general, the initial mass will be reduced for cases in which the total plant mass is less than the propellant mass it will produce over time. Because of the exponential decreases in mass involved in orbital mechanics, ISRU may reduce the initial spacecraft mass by thousands of kilograms for a typical sample return mission to Mars. It should be noted that although the use of ISRU is advantageous in a substantial majority of the studied cases, ISRU will not provide benefits in every conceivable scenario. The extent of its usefulness is highly dependent upon the propellant combination, the return payload, and the production plant 114 ProDellant Combination FoM R-factor H2 + 02 0.159 0.6 CH30H + 02 0.369 0.5 CH4 + 02 0.319 0.7 NH3 + 02 0.167 0.5 C2H50H + 02 0.174 0.5 CO + 02 0.019 0.4 H2 + N204 0.050 0.4 CH4 + N204 0.358 0.3 CO + N204 0.021 0.5 C2H50H + N204 0.142 0.3 NH3+ N204 0.016 0.3 Table 17 - R-factor Comparison mass. If a very small mass is to be returned from Mars, ISRU would only benefit the mission by proving the feasibility and reliability of the technology. Figure 32 shows the FoM as a function of the payload mass returned to LEO for both conventional and ISRU sample return missions. The application of ISRU is beneficial for the methane and oxygen propellant combination when a sample mass greater than 1 kg is to be returned. It can be observed from the results in Figure 32 that in general, the greater the mass to be returned, the more beneficial ISRU will be. To determine the extent of the benefits of ISRU on manned missions, significantly larger payload return masses were input into the FoM spreadsheet program to calculate the mass savings for much larger vehicles. Return payload masses were examined ranging from 2 to 10,000 kg. From the graph in Figure 33, it should be noted that the FoM for a specific propellant combination will vary in a unique manner for payload masses across this large payload range. The methane and oxygen combination using ISRU is the most advantageous 115 Propellant Combination • CH4 + 02 0.7 E222 All Earth Carrlad E23 I8RU (R-factor • 0.7) (R-la 1.0 1.5 2.0 2.6 Sample mass returned to LEO, kg Figure 32: Sample Mass Variations for sample return missions, but the benefits rapidly level off with increasing payloads. Using hydrogen fuel for the outbound legs and methane fuel for the inbound legs provides the highest FoM for the larger payload masses. However, the benefits appear to be leveling off for the largest payload using this joint combination while the strict hydrogen and oxygen case using ISRU appears to be increasing steadily in FoM. Over this entire payload range, it should be emphasized that the ISRU mission cases provide the highest comparative FoM values. Although the propellant processed through the use of ISRU may not always yield optimum performance in comparison to other propellant combinations, the benefits to be 116 MARS MISSION VARIATIONS Figure-of-Merit vs. Returned Mass 8 H2+02 (All-Earth) |R-0.6| • H2+02 (ISRU) IR-0.5I o B CH4+02 (ISRU) (R-0.61 u_ H2+02 out/CH4»02 In (R-0.61 ® 2 M— o £> k. 3 O) 100 1000 10000 Mass returned to LEO, kg LEO * 500 km altitude Figure 33: Return Payload Variations gained through the use of ISRU may outweigh any disadvantages in terms of overall mission planning. Accurate comparison studies must made in order to determine which mission strategies will be most influenced by ISRU, and which mission profiles stand to gain the largest benefits. Figure 34 presents the calculated FoM values when using different fuel combinations in the studied MSR mission. All of the results in the figure use liquid oxygen as the oxidizer for the complete mission. For this comparison study, the mass of the oxygen plant is fixed at 100 kg and a sample mass of 2 kg is to be returned to LEO. The propellant combinations shown in the figure are indicative of the results calculated for all of the analyzed 117 propellant combinations. A significant number of cases showed a moderate increase in the FoM when ISRU is incorporated into the mission plan. Once again, even a moderate increase in the FoM value should be capitalized upon since it translates into a mass savings on the order of thousands of kilograms in the initial mass of the spacecraft. A few rare cases showed a nominal decrease in the FoM when using ISRU, but this trend quickly reverses itself when using larger payload masses. Most importantly, the use of ISRU in certain mission cases induced a remarkable increase in the FoM. For example, Figure 34 indicates that the initial mass is reduced by 15,453 kg when ISRU is used for the methane and oxygen propellant combination, resulting in the highest FoM. For this case, the use of ISRU has Mars Sample Return Mass - 2 kg Oxidizer - 02 0.6 | All Earth Carried E33 isru (R-fsctor A* -0.1) YSSAWA H2 CH30H CH4 NH3 C2HSOH CO Fuel Choice Figure 34: ISRU Comparison 118 resulted in an FoM increase of 13% and an initial mass reduction of more than 25%. The relative changes in the FoM and the initial mass are not equal because the ISRU case has a greater risk associated with the new technology, and the FoM calculation factors in this risk assessment. If the original propellant plant can sustain extended production times, a fuel depot could be established on the Martian surface which would substantially reduce the payloads and launch masses of all later missions. The argument that certain propellant combinations should use ISRU is compelling, but the actual implementation of this strategy must prove to be both practical and reliable. The complexity of the production plant, the reliability of its components, and the mass of the total system will be the driving design factors. The hardware should be kept as simple and light-weight as possible. Figure 35 details the relative benefits of decreasing the plant mass. The total plant mass was initially projected to be 197 kg, but the technological advances that have occurred since the inception of the oxygen plant project have lowered the predicted mass to approximately 100 kg. A reduction of 97 kg in the oxygen plant mass translates into a spacecraft initial mass savings of 8046 kg and an FoM increase of 0.0555. The necessary production rate will dictate the total system mass, and as a general rule, the system mass will be a function of the production rate raised to the two-thirds power. The power source will ultimately account for a large percentage of the total production plant mass. Therefore, further effort should be focused on the reduction of both the plant power requirements and the specific weight of the power source. Another technique to increase the FoM is to optimize the vacuum specific impulse. There are several ways to alter the specific impulse, but the greatest impact is caused by variations in the fuel and oxidizer selection. The problem lies in determining if the optimum 119 Mars Sample Return Mass - 1 kg Propellant Combination - CH4 + 02 0.40 60 —r 0.35 ~ 0.30 0) 2 JL ? 0.25 £ J? E 0.20 FOM ™f™ M(LEO) 0.15 20 50 100 150 200 250 300 350 Oxygen Production Plant Mass, kg Figure 35: Plant Mass Variations engine design parameters also maximize the resultant FoM. Graphical design solutions are presented in Figures 36 and 37 for a methane and liquid oxygen propellant combination with ISRU. This combination is of particular interest because of its promising FoM with respect to all of the permutations studied. The maximum FoM occurs for an oxidizer to fuel mass ratio that also results in a maximum vacuum specific impulse. For this propellant combination, the optimum O/F ratio is 2.8 (stoichiometric = 4.0). However, the maximum FoM does not occur for a nozzle area ratio that maximizes the vacuum specific impulse. While an increase in the nozzle area ratio will always result in an increase in the specific impulse, there is a transition point at which the additional mass of the nozzle will offset the 120 Propellant Combination - CH4 + 02 (ISRU) Mars sample return mass - 1 kg 0.35 400 0.30 FoM "+••• Isp(vac) 2 300 £ 0.25 0.20 - 200 o• 0.16 © 0.10 u. 100 0.06 0.00 0 2 4 6 8 10 12 14 16 Oxidizer/Fuel Ratio (O/F) Figure 36: Fuel Ratio Effects gains made by the increase in specific impulse. This transition point can only be determined from an examination of the overall spacecraft design. The data presented in all of the performance plots are based upon a nozzle area ratio of SO. An initial was made that this value would be closest to the optimum nozzle area ratio. Since the publication of these graphs,51 however, it has been determined that the optimum nozzle area ratio is 25. Consequently, if all of the previous data are fully optimized with the appropriate nozzle area ratio, the FoM would increase by approximately 15% and the initial masses would decrease by a corresponding 15%. Similar optimization plots can also be generated for conventional- missions without ISRU. 121 Propellant Combination - CH4 • 02 (ISRU) Mars sample return mass - 1 kg 0.40 400 0.35 o & 0.30 360 o © CO o> 0.25 -o 01 2 0.20 300 £ 3 cn Li. 0.16 FOM Isp(vac) 0.10 260 20 40 60 80 100 120 Nozzle Area Ratio (Ae/At) Figure 37: Nozzle Area Ratio Effects Figure 38 shows the increase in the FoM that results from the use of modular engines. A methane and liquid oxygen propellant combination is again presented for analysis which will return a 2 kg soil sample mass. The conventional mission with all Earth carried propellants is shown along with the ISRU case. Superimposed on the plot are the FoM's if the R-factor is set equal to unity. In an absolute sense, the FoM values for the missions using modular engines decrease because the initial mass increases slightly. Although the resultant masses are greater, the FoM has a relative increase after the appropriate R-factors are included in the calculation. The risk has decreased as a result of increases in the reliability and the reparability. For missions utilizing modular engines, the R-factor is increased by 0.1. 122 Conversely, the plot shows that the inclusion of ISRU will decrease the R-factor by 0.1. The initial mass savings are so great, however, that the mission design is still highly favorable Propellant Combination - CH4 + 02 Mars sample mass returned • 2 kg i " "|" 1 'T' I ALL EARTH MODULAR I8RU I8RU/MODULAR CARRIED Mission Configurations Figure 38: Technology Benefits despite the added risks. The identical results are also presented in Figure 39 showing the initial masses in LEO for the four distinct mission cases. The R-factor has no effect on the mass values, so the ISRU plan without a modular engine design provides the lowest initial mass. If this mission plan is selected, the initial mass required in LEO to successfully complete the mission would be approximately 43,000 kg. Table 18 summarizes the relatives changes for the mission variations using the conventional mission for comparison. Both the changes in the initial mass at LEO and the FoM values are presented. 123 Propellant Combination • CH4 • 02 Mars sample mass returned - 2 kg 70 61.000 60 50 a 45,100 oo 42.800 40 2 30 - o hi 20 - 2 10 - 0 ALL EARTH MODULAR I8RU I8RU/MODULAR CARRIED Mission Configurations Figure 39: Technology Mass Savings Mission Variation AM (LEO) AFoM All Earth Carried - - Modular Engines +6.4% +8.9% ISRU -25.3% +12.7% ISRU/Modular Engines -21.2% +28.8% Table 18 - Mission Variation Summary CHAPTER 12 PROJECT SUMMARY Three fundamental objectives have been identified and achieved in this thesis. The first objective was to identify new technologies which have the potential to improve future space missions. ISRU and modular engines were shown to be promising technologies. The second goal was to provide a detailed user's manual for the FoM spreadsheet program. The FoM program contains hundreds of complex equations which use a multitude of interdependent variables. This thesis functions as a manual which defines the initial approximations and limitations of the equations used in the FoM calculations. The third objective was to demonstrate the usefulness of the FoM program by using it to assess and compare several Mars Sample Return mission plans. Throughout the preliminary mission design process, the program provided a great deal of information for improving the efficiency of specific mission plans. By adjusting certain parameters, the optimum design points were easily determined through an integrated use of numerical and graphical data analysis. The Figure-of-Merit approach provides many benefits which are not available from other analysis techniques. Most important, the FoM yields quantitative results, which provide hard, complete data necessary for planning decisions in the space program. Second, the method focuses on overall mission design rather than on specific component optimization. Individual components will be optimized only to the point where it starts to adversely affect the mission. By viewing this design problem in a global perspective, the tradeoff point will become more obvious. Third, the FoM strategy offers a uniform method of making side-by- side comparisons of similar missions. For mission plans to be accurately compared, the same 125 analysis procedure must be used for each case. Finally, the FoM itself is an effective screening parameter for a wide variety of missions. Numerous mission modifications can be easily made within the spreadsheet format, and the resultant FoM will indicate whether the change was advantageous or not. In the current scheme all FoM definitions are modified by the R-factor, which accounts for the risk, reliability, and reparability factors. A relative approach for the R-factor is used, and every effort is made to include accurate data despite the subjective nature of the factors. While the absolute R-factor may not be exactly correct for a given mission, the relative effects are quite acceptable for comparison purposes. The FoM spreadsheet has been used to analyze a MSR mission. The three most promising propellant combinations are methanol/oxygen, methane/nitrogen tetroxide, and methane/oxygen. Any one of these three combinations would be an excellent choice over any of the other combinations studied. The methane/oxygen case was ultimately selected as the most promising propellant combination because both the fuel and the oxidizer can be produced on the surface of Mars by ISRU. The chemical reactions and processes have been studied in detail and are currently the focus of concentrated research efforts. Historically, methanol has been a more common fuel choice than methane, and it too could be produced on the surface of Mars. The analysis indicates that methanol is an excellent fuel choice when combined with oxygen, and it is recommended that future research be focused on the ISRU production of methanol. However, the methanol/oxygen combination has a lower FoM for sample class payloads than the methane/oxygen combination in missions with ISRU production of oxygen. The use of nitrogen tetroxide as an oxidizer reduces the rocket engine performance, but may be advantageous to the overall system because it does not require any special refrigeration. The methane/nitrogen tetroxide propellant combination requires a more 126 complete analysis to fully assess its potential advantages over the chosen methane/liquid oxygen combination. The results indicate that the most effective mission plan uses a methane/liquid oxygen propellant combination incorporating both ISRU and modular engines. The FoM program indicates that the optimum engine design for this propellant combination should use an oxidizer to fuel mass ratio of 2.8, a chamber pressure of 500 psia, and a nozzle area ratio of 25. Using these design parameters, a 2 kg Mars sample could be returned to LEO by a spacecraft with an initial mass of approximately 45,200 kg in LEO. The FoM for this configuration is 0.5227. The payload includes an oxygen production plant which weighs 100 kg and produces 10 kg/day of oxygen on the surface of Mars. For a MSR mission, the production rate could be reduced to 1.1 kg/day and still provide an adequate margin of safety. With this lower production rate, the initial spacecraft mass in LEO would be reduced by 15%, resulting in a 15% increase in FoM. When the plant mass was reduced by 49% from the original oxygen plant mass estimates, the FoM increased by 17% while the initial spacecraft mass in LEO is decreased by 17%. The use of modular engines only will increase the FoM by 9%, while the use of ISRU only will increase the FoM by 13%. Combining both technologies, the FoM can be increased by a full 29%. This translates into an initial spacecraft mass savings of 21% in LEO. For manned missions, or for any mission that requires the return of greater payload masses, methane is a poor fuel choice for the outbound journey. It is more advantageous to use the most energetic propellant combination of liquid hydrogen/liquid oxygen for the outbound journey and later convert to the use of methane/oxygen for the return voyage. From the analysis, it is clear that one propellant combination will not provide optimum results 127 for every conceivable space mission. Different missions have different requirements, and each specific mission must be analyzed in an overall system perspective to obtain the optimum results. The quantitative FoM approach is a concise and effective planning tool for the analysis and design of future space missions. APPENDIX A FIGURE-OF-MERIT SPREADSHEET 129 A B C D E F 1 2 MARS SAMPLE RETURN MISSION ANALYSIS Bruce Preiss 4/09/91 3 4 (Total sheets - 13, review sheet 13 for map and additional info.) 5 6 7 SHEET #1 - PROGRAM SUMMARY 8 9 MASS SUMMARY FOR EACH MISSION 10 11 Case 1 Case 2 Case 3 Case 4 12 BBBBBBBB BBBBBBBB BBBBBBBB BBBBBBBB 13 FUEL (outbound) H2 H2 CH4 H2 14 OXIDIZER (outbnd) 02 02 02 02 15 ISPP FUEL 16 ISPP OXID. 02 02 02 17 FUEL (return) H2 H2 CH4 CH4 18 OXIDIZER (return) 02 02 02 02 19 ENGINES (return) Conventnl Conventnl Conventnl Conventnl 20 21 Case 1 Case 2 Case 3 Case 4 22 23 MA [kg] 148 140 66 101 24 MF [kg] 5976 5724 6474 4449 25 MFL [kg) 450 430 23 325 26 MFT [kg] 1115 1071 212 836 27 MG [kg] 4126 3962 2272 3114 28 MFS [kg] 5708 5480 3131 4302 29 ME [kg] 1822 1749 1001 1374 30 MN [kg] 4033 3872 2222 3042 31 MNYFAC [kg] 0 0 0 0 32 MOX [kg] 33464 31997 18053 24766 33 MOXFAC [kg] 0 100 100 100 34 MOXL [kg] 252 240 126 182 35 MOXT [kg] 375 361 219 287 36 MRF [kg] 5417 5268 0 4179 37 MROV [kg] 150 150 150 150 38 MROX [kg] 2853 2659 1797 2229 39 MS [kg] 3979 3821 2187 3001 40 MSAMPL [kg] 1 1 1 1 41 MSUPP [kg] 1 1 1 1 42 43 MLEO [kg] 69168 66356 37885 51933 44 45 RF=INV RISK 0.40 0.30 0.59 0.57 46 47 FoM 0.1157 0.0904 0.3115 0.2195 48 49 Case 1 Case 2 Case 3 Case 4 130 BABC D £ p Q 1 SHEET #2 - MISSION DETAILS 2 3 PHYSICAL CONSTANTS 4 5 RADIAL DISTANCE FROM BODY TO SUN, EARTH-SUN REFERENCED 6 EARTH - 1 AU MARS » 1.524 AU 7 8 MEAN RADIAL DISTANCE FROM BODY TO SUN [m] 9 EARTH - 1.50E+11 m MARS - 2.28E+11 m 10 11 GRAVITATIONAL ACCELERATION ON THE BODY SURFACE 12 GZERO » 9.8100 m/s2 GOMARS 3.7300 m/s2 13 14 15 PROPULSIVE VELOCITY REQUIREMENTS m/s 16 17 LEG 1 LEO to HOHMANN towards Mars 3800 18 (No MID-COURSE CORRECTION used) 0 19 Number of stages = 2 20 Aerobraking = No 21 Modular engines «= No 22 2 3 LEG 2 HOHMANN to Mars Parking Orbit (200 km) 2700 24 (includes a 0.5 [km/s] MID-COURSE CORR.) 500 25 Number of stages • 2 26 Aerobraking « No 27 Modular engines = No 28 29 LEG 3 MPO to Mars Soft Landing 1000 30 (No MID-COURSE CORRECTION used) 0 31 Number of stages * 1 32 Aerobraking •= Yes 33 . Modular engines = No 34 35 LEG 4 Mars Ascent Vehicle to MPO 3500 36 (No MID-COURSE CORRECTION used) 0 37 Number of stages = 2 38 Aerobraking » No 39 Modular engines » No 40 41 LEG 5 MPO to HOHMANN towards Earth 2700 42 (No MID-COURSE CORRECTION used) 0 4 3 Number of stages • 2 44 Aerobraking • No 45 Modular engines - No 46 47 LEG 6 HOHMANN to Highly eccentric Earth orbit 1000 48 (includes 0.5 [km/sJ MID-COURSE CORR.) 500 49 Number of stages * 1 50 Aerobraking • Yes 51 Modular engines = No 52 53 LEG 7 N/A 0 131 BAB C D EF G 54 (No MID-COURSE CORRECTION used) 0 55 Number of stages - 0 56 Aerobraklng « No 57 Modular engines - No 5a 59 LEG 8 N/A 0 60 (No MID-COURSE CORRECTION used) 0 61 Number of stages • 0 62 Aerobraking • No 63 Modular engines • No 64 65 TOTAL DELTA V 15700 66 132 ABC D E 1 SHEET #3 - ROCKET PERFORMANCE DATA 2 Case 1 ALL EARTH TRANSPORTED LH2 AND LOX 3 4 [Refrig.] 5 Fuel (Outbound) H2 Yes 6 Oxidizer(Outbound) 02 YeB 7 Oxidizer/Fuel Ratio 5.6 kg OX/kg fuel 8 Chamber Pressure 500 pBia 9 Nozzle Expansion Ratio 50 10 Specific Impulse 445.95 sec 11 Characteristic Velocity 7555 ft/sec 12 13 (Refrig.] (Surf. Refrig.] [ISRU] 14 Fuel (Return) H2 No Yes No 15 Oxidizer (Return) 02 No Yes No 16 Oxidizer/Fuel Ratio 5.6 kg OX/kg fuel 17 Chamber Pressure 500 pBia 18 Nozzle Expansion Ratio 50 19 Specific Impulse 445.95 sec 20 Characteristic Velocity 7555 ft/sec 21 22 Case 2 ALL EARTH TRANSPORTED LH2 AND ISRU LOX 23 24 [Refrig. 25 Fuel (Outbound) H2 Yes 26 Oxidizer(Outbound) 02 Yes 27 Oxidizer/Fuel Ratio 5.6 kg OX/kg fuel 28 Chamber Pressure 500 psia 29 Nozzle Expansion Ratio 50 30 Specific Impulse 445.95 sec 31 Characteristic Velocity 7555 ft/sec 32 33 [Refrig.] [Surf. Refrig.] [ISRU] 34 Fuel (Return) H2 NO Yes No 35 Oxidizer (Return) 02 No No Yes 36 Oxidizer/Fuel Ratio 5.6 kg OX/kg fuel 37 Chamber Pressure 500 psia 38 Nozzle Expansion Ratio 50 39 specific Impulse 445.95 sec 40 Characteristic Velocity 7555 ft/sec 133 CABCD E F G 41 42 Case 3 ALL EARTH TRANSPORTED CH4 AND ISRU LOX 43 4 4 [Refrlg.] 45 Fuel (Outbound) CH4 No 46 Oxidlzer(Outbound) 02 Yes 47 Oxidizer/Fuel Ratio 2.8 kg OX/kg fuel 48 Chamber Pressure 500 psla 49 Nozzle Expansion Ratio 50 50 Specific Impulse 356.05 sec 51 Characteristic Velocity 6047 ft/sec 52 53 [Refrig.] [Surf. Refrig.] [ISRU] 54 Fuel (Return) CH4 No No No 55 oxidizer (Return) 02 No No Yes 56 Oxidizer/Fuel Ratio 2.8 kg OX/kg fuel 57 Chamber Pressure 500 psia 58 Nozzle Expansion Ratio 50 59 Specific Impulse 356.05 sec 60 Characteristic Velocity 6047 ft/sec 61 62 Case 4 LH2+LOX (out)/EARTH TRANSP. CH4+ISRU LOX (in) 63 64 [Refrig.] 65 Fuel (Outbound) H2 Yes 66 Oxidizer(Outbound) 02 Yes 67 Oxidizer/Fuel Ratio 5.6 kg OX/kg fuel 68 Chamber Pressure 500 psia 69 Nozzle Expansion Ratio 50 70 Specific Impulse 445.95 sec 71 Characteristic Velocity 7555 ft/sec 72 73 [Refrig.] [Surf. Refrig.] [ISRU] 74 Fuel (Return) CH4 No No No 75 Oxidizer (Return) 02 No No Yes 76 Oxidizer/Fuel Ratio 2.8 kg OX/kg fuel 77 Chamber Pressure 500 psia 78 Nozzle Expansion Ratio 50 79 Specific Impulse 356.05 sec 80 Characteristic Velocity 6047 ft/sec 134 I J K L M N 1 DATABASE RETRIEVAL INFORMATION 2 Field list for paradox file PROPEL.DB converted to 3 dBase III format in file PROPELDB.DBF 4 5 6 FUEL Character 4 NA 7 OXIDIZER Character 5 NA 8 OFRATIO Numeric 19,4 NA 9 PC Numeric 19,4 NA 10 TC Numeric 19,4 NA 11 ARATIO Numeric 19,4 NA 12 PRATIO Numeric 19,4 NA 13 CSTAR Numeric 19,4 NA 14 ISP Numeric 19,4 NA 15 ISPVAC Numeric 19,4 NA 16 17 18 Criteria and Output ranges are below for database queri 19 20 21 Criteria Range 22 23 FUEL OXIDIZER OFRATIO PC TC ARATIO 24 500 50 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 Output Range 42 43 FUEL OXIDIZER OFRATIO PC TC ARATIO 44 CH30H N204 3.1266 500 3183.2 50 45 CH30H N204 7.2954 500 2339.1 50 46 CH30H N204 10.422 500 1879.7 50 47 CH30H N204 15.633 500 1415.7 50 48 CH30H N204 26.055 500 962.9 50 49 CH30H N204 36.477 500 740.4 50 50 135 c 1 2 3 4 5 6 NA NA 7 NA NA 8 NA NA 9 NA NA 10 NA NA 11 NA NA 12 NA NA 13 NA NA 14 NA NA 15 NA NA IS 17 18 19 20 21 22 23 PRATIO CSTAR ISP ISPVAC 24 25 26 27 26 29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 PRATIO CSTAR ISP ISPVAC 44 668.36 5124 293.05 305.25 45 885.29 4108.5 227.55 234.75 46 1009.66 3623.5 198.45 204 47 1108.68 3100.5 167.6 171.95 48 403.1 1260.5 67.4 69.8 49 466.61 1096 59.3 61.15 50 136 DA B CD E F 6 H 1 SHEET #4 - ROCKET PROPELLANT DATA 2 3 FUEL AND OXIDIZER COMBINATIONS AVAILABLE IN PARADOX DATABASE 4 5 Stoichiometric Oxidizer/Fuel Ratio 6 Earth region 7 H2 + 02 Liquid hydrogen 8.000 8 N2H4 + N204 Hydrazine 1.438 9 NH3 + N204 Anunonia+Nitrogen tetroxlde 2.029 10 H2 + H202 LH2 + Hydrogen peroxide 17.000 21 12 Moon region 13 CH4 + 02 Methane 4.000 14 SiH4 + 02 Silane 2.000 15 A1 + 02 Aluminum powder 1.778 16 H2 + A1 + 02 lot A1 7.289 17 H2 + A1 + 02 20% A1 6.578 18 C6H802(OH)2 + 02 Nylon 1.908 19 20 Mars region 21 N2H4 + 02 1.000 22 NH3 + 02 1.627 23 CH4 + 02 4.000 24 CO + 02 Carbon monoxide 1.143 25 CO + H202 1.213 26 CO + N204 0.821 27 C2H50H + 02 Ethyl alcohol 2.0B7 28 CH30H + 02 Methyl alcohol 1.500 29 CH30H + H202 3.18B 30 CH4 + H202 8.500 31 CH4 + N204 32 NH3 + H202 33 N2H4 + H202 34 CH30H + N204 35 C2H50H + N204 36 C2H50H + H202 37 38 39 DATABASE PARAMETERS 40 41 Chamber Pressures (psia) 42 Earth region 50, 100, 500, 1000, 2000 43 Moon region 0.1, 10, 50, 100 44 Mars region 1, 10, 50, 100, 500 45 46 Nozzle Area Ratios 47 All regions 5, 10, 25, 50, 100, 250 48 49 Oxidizer/Fuel Ratios 50 All regions use stoichiometric values given above with 51 multipliers of: 0.3, 0.7, 1.0, 1.5, 2.5, 3.5 52 137 DA B CD E F G H 54 PROPELLANT SPECIFIC VOLUME (at a given tank pressure) 55 - - 56 (Saturated liquid values at varying pressures) 57 (Represents worst case values in calculations) 58 59 Pc (psi) Ptank (MP 61 Note: Assume tank 1.00 0.010342 62 pressure is 10.00 0.103421 63 1.5 times 33.33 0.344737 64 greater than 50.00 0.517105 65 the chamber 100.00 1.034210 66 pressure. 500.00 5.171050 67 68 Table of Specific Volumes (mA3/kg) 5 9 BSD8BS80BBIBBBCBIBBRH9BBBBBBIBIBBBB 70 Tank Pressures in MPa 71 0.010342 0.103421 0.344737 0.517105 1.034210 72 Chemical species 73 H2 0.013076 0.014142 0.015592 0.016552 0.022494 74 N2H4 0.000989 75 NH3 0.001377 0.001468 0.001547 0.001583 0.001662 76 02 0.000816 0.000878 0.000934 0.000962 0.001028 77 N204 0.000687 0.000690 0.000749 0.00078 0.000812 78 H202 0.000691 79 CH4 0.002215 0.002368 0.002524 0.002601 0.002790 80 SiH4 81 A1 82 H2 + A1 83 H2 + A1 84 C6H802(OH)2 85 CO 0.001182 0.001268 0.001364 0.001417 0.001541 86 C2H50H 0.001250 0.001323 0.001392 0.001427 0.001512 87 CH30H 0.001263 0.001332 0.001395 0.001432 0.001502 138 D I 54 55 56 57 58 59 a) 60 61 62 63 64 65 66 67 68 69 70 71 5. 171050 72 73 C.031888 74 75 0.002076 76 0.002293 77 0.001061 78 79 0.006233 80 81 82 83 84 85 0.003322 86 0.002126 87 01.00197 139 D A B C D E 89 Table of Storage Temperatures (deg K) 90 91 Tank Pressures in MPa 92 0.010342 0.103421 0.344737 0.517105 1.034210 93 Chemical Species 94 H2 14.44 20.35 25.22 27.29 31.44 95 N2H4 96 NH3 202.27 240.20 267.40 278.21 299.14 97 02 72.47 90.35 103.64 109.20 120.19 98 N204 263.91 291.69 322.25 333.36 347.25 99 H202 100 CH4 90.6B 111.86 128.82 135.85 149.84 101 S1H4 102 A1 103 H2 + A1 104 H2 + A1 105 C6H802(OH)2 106 CO 68.14 81.79 94.03 99.24 109.48 107 C2H50H 302.49 352.26 385.91 399.40 425.36 108 CH30H 293.24 338.25 370.48 386.21 411.77 109 110 PTank Row Column 111 Case 1 112 Fuel • • Out H2 0.344737 113 Ox. - Out 02 0.344737 114 Fuel - • In H2 5.17105 115 Ox. - In 02 5.17105 116 Case 2 117 Fuel - Out H2 0.344737 118 Ox. - Out 02 0.344737 119 Fuel • In H2 5.17105 120 Ox. - In 02 5.17105 121 Case 3 122 Fuel • Out CH4 0.344737 123 Ox. - Out 02 0.344737 124 Fuel • In CH4 5.17105 125 Ox. - In 02 5.17105 126 Case 4 127 Fuel • Out H2 0.344737 128 Ox. - Out 02 0.344737 129 Fuel • In CH4 5.17105 130 Ox. - In 02 5.17105 D I 89 90 91 92 5. 171050 93 94 32.94 95 96 363.64 97 154.58 98 405.59 99 100 190.55 101 102 103 104 105 106 132.92 107 503.94 108 488.20 109 110 Spec. Vol 111X 1X 112 0. 031888 113 0. 002293 114 0. 015592 115 0.000934 116 117 0.031888 118 0. 002293 119 0. 015592 120 0.000934 121 122 0. 006233 123 0. 002293 124 0. 002524 125 0. 000934 126 127 0. 031888 128 0. 002293 129 0. 002524 130 0. 000934 141 A B CD SHEET #5 - ISRU COMPONENT DATA IN-SITU RESOURCE UTILIZATION SCALING EQUATIONS 1 5 CASE 1 MASSES REQUIRED 6 PROD. RATE 7 PRODUCTION RATE (kg 02/dav) 10 0:3 8 PRODUCTION TIME (days) 200 9 POWER SUPPLY SP. WT. (kg/kwe) 30 10 11 FACTORY COMPONENTS Mass (kg) Watts-electric Watts-thermal 12 FILTER 0,.46 4.92 13 C02 ACCUMULATOR 12,.30 330.00 14 HEAT EXCHANGER 4..89 208.30 15 ZIRCONIA CELLS 17..50 849.94 3206.00 16 RADIATOR 2,.26 17 MEMBRANE SEPARATOR 1,.36 18 RECYCLE COMPRESSOR 20..45 199.75 19 02 COMPRESSOR 31..50 945.75 20 INT. STORAGE 5..36 21 PIPING / TUBING 6..00 22 VALVES 31.60 23 COMPUTER + COMM. 10..21 24 MASS MARGIN 5..61 25 26 SUBTOTALS 117..89 1416.22 4360.05 27 28 HEAT SOURCE 36..62 29 POWER SOURCE 42..49 30 31 TOTAL MASS 100.0024 > 32 SSSSBBBSSS 3EB3BBCCCBBS3 BIBBBBBBIBBSBRSIHIItSS 33 34 35 36 CASE 2 MASSES REQUIRED 37 PROD. RATE 38 PRODUCTION RATE (kg 02/day) 10 0.3 39 PRODUCTION TIME (days) 200 40 POWER SUPPLY SP. WT. (kg/kwe) 30 41 42 FACTORY COMPONENTS Mass (kg) Watts-electric Watts-thermal 43 FILTER 0.46 4.92 44 C02 ACCUMULATOR 12.30 330.00 45 HEAT EXCHANGER 4.89 208.30 46 ZIRCONIA CELLS 17.50 849.94 3206.00 47 RADIATOR 2.26 48 MEMBRANE SEPARATOR 1.36 49 RECYCLE COMPRESSOR 20.45 199.75 50 02 COMPRESSOR 31.50 945.75 142 B C D E 51 INT. STORAGE 5.36 52 PIPING / TUBING 6.00 53 VALVES 31.60 54 COMPUTER + COMM. 10.21 55 MASS MARGIN 5.61 56 57 SUBTOTALS 117.89 1416.22 4360.05 58 59 HEAT SOURCE 36.62 60 POWER SOURCE 42.49 61 62 . MASS 100.0024 > CASE 2 63 sssessss 64 65 66 67 CASE 3 MASSES REQUIRED 68 PROD. RATE 69 PRODUCTION RATE (kg 02/day) 10 0.4 70 PRODUCTION TIME (days) 200 71 POWER SUPPLY SP. WT. (kg/kwe) 30 72 73 3RY COMPONENTS Mass (kg) watts-electric Watts-thermal 74 FILTER 0 .46 4 .92 75 C02 ACCUMULATOR 12 .30 330 .00 76 HEAT EXCHANGER 4 .89 208.30 77 ZIRCONIA CELLS 17 .50 849..94 3206.00 78 RADIATOR 2 .26 79 MEMBRANE SEPARATOR 1,.36 80 RECYCLE COMPRESSOR 20..45 199..75 81 02 COMPRESSOR 31,.50 945.75 82 INT. STORAGE 5,,36 83 PIPING / TUBING 6,,00 84 VALVES 31..60 85 COMPUTER + COMM. 10.21 86 MASS MARGIN 5,.61 87 88 SUBTOTALS 117,.89 1416..22 4360.05 89 90 HEAT SOURCE 36.,62 91 POWER SOURCE 42..49 92 93 . MASS 100.0024 > CASE 3 94 95 96 97 98 CASE 4 MASSES REQUIRED 99 PROD. RATE PRODUCTION RATE (kg 02/day) 10 0.4 143 E B C D E 101 PRODUCTION TIME (days) 200 102 POWER SUPPLY SP. WT. (kg/kwe) 30 103 104 )RY COMPONENTS Mass (kg) Watts-electric Watts -thermal 105 FILTER 0.46 4.92 106 C02 ACCUMULATOR 12.30 330.00 107 HEAT EXCHANGER 4.89 208.30 108 ZIRCONIA CELLS 17.50 849.94 3206.00 109 RADIATOR 2.26 110 MEMBRANE SEPARATOR 1.36 111 RECYCLE COMPRESSOR 20.45 199.75 112 02 COMPRESSOR 31.50 945.75 113 INT. STORAGE 5.36 114 PIPING / TUBING 6.00 115 VALVES 31.60 116 COMPUTER + COMM. 10.21 117 MASS MARGIN 5.61 118 119 SUBTOTALS 117.89 1416.22 4360.05 120 121 HEAT SOURCE 36.62 122 POWER SOURCE 42.49 123 124 . MASS 100.0024 > CASE 4 125 '88 126 144 A B C D l SHEET #6 - SUPPORT COMPONENTS 2 3 4 REFRIGERATION REQUIREMENTS FOR PROPELLANT WHILE ON THE MARTIAN 5 SURFACE (BASED ON RADIATIVE AND FREE-CONVECTIVE HEAT FLUX) 6 7 ATMOSPHERIC DENSITY [kg/m3] 0.1221001 8 AMBIENT T (PEAK) [K] 260.00 9 VISCOSITY [kg/ms] 0.0000131 10 ISOBARIC COMP [1/K] 0.0038462 11 PRANDTL NUMBER 0.76400 12 THERMAL CONDUCTIVITY [W/mK] 0.01385 13 14 DIAMETERS OF TANKS (FOUR PER STAGE) 15 Case 1- —Case 2-- 16 OXIDIZER FUEL OXIDIZER FUEL 17 DLEG6IV [m] O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 18 DLEG6III [m] 0.00E+00 O.OOE+OO O.OOE+OO O.OOE+OO 19 DLEG6II [m] O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 20 DLEG6I [m] 1.88E-01 2.55E-01 1.88E-01 2.55E-01 21 DLEG5IV [m] O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 22 DLEG5III [m] 0.00E+00 O.OOE+OO O.OOE+OO O.OOE+OO 23 DLEG5II [m] 2.23E-01 3.02E-01 2.23E-01 3.02E-01 24 DLEG5I (ni J 2.73E-01 3.69E-01 2.73E-01 3.69E-01 25 DLEG4IV [m] O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 26 DLEG4III [m] O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 27 DLEG4II [m] 3.98E-01 5.39E-01 3.98E-01 5.39E-01 28 DLEG4I [m] 5.41E-01 7.33E-01 5.41E-01 7.33E-01 29 30 GRASHOFF NUMBERS 31 Case 1— Case 2-- 32 OXIDIZER FUEL OXIDIZER FUEL 33 GR6IV O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 34 GR6III O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 35 GR6II O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 36 GR6I 1.34E+06 4.94E+06 1.34E+06 4.94E+06 37 GR5IV O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 38 GR5III O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 39 GR5II 2.24E+06 8.25E+06 2.24E+06 8.25E+06 40 GR5I 4.08E+06 1.51E+07 4.08E+06 1.51E+07 41 GR4IV O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 42 GR4III O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 43 GR4II 1.27E+07 4.69E+07 1.27E+07 4.69E+07 44 GR4I 3.19E+07 1.18E+08 3.19E+07 1.18E+08 45 145 F ABCD E F G 46 RAYLEIGH NUMBERS 47 Case 1-- Case 2— 48 OXIDIZER FUEL OXIDIZER FUEL 49 RA6IV 0.00E+00 O.OOE+OO O.OOE+OO 0.OOE+OO 50 RA6III O.OOE+OO O.OOE+OO O.OOE+OO 0.OOE+OO 51 RA6II 0.00E+00 O.OOE+OO O.OOE+OO 0.OOE+OO 52 RA6I 1.02E+06 3.77E+06 1.02E+06 3 •77E+06 53 RA5IV O.OOE+OO O.OOE+OO O.OOE+OO 0.OOE+OO 54 RA5III O.OOE+OO O.OOE+OO O.OOE+OO 0.OOE+OO 55 RA5II 1.71E+06 6.31E+06 1.71E+06 6.31E+06 56 RA5I 3.12E+06 1.15E+07 3.12E+06 1. 15E+07 57 RA4IV O.OOE+OO O.OOE+OO O.OOE+OO 0.OOE+OO 58 RA4III O.OOE+OO O.OOE+OO O.OOE+OO 0.OOE+OO 59 RA4II 9.70E+06 3.58E+07 9.70E+06 3.58E+07 60 RA4I 2.44E+07 8.99E+07 2.44E+07 8.99E+07 DA 1 62 NUSSELT NUMBERS 63 Case 1-- Case 2-- 64 OXIDIZER FUEL OXIDIZER FUEL 65 NU6IV 2.OOE+OO 2.OOE+OO 2.OOE+OO 2.OOE+OO 66 NU6III 2.00E+00 2.OOE+OO 2.OOE+OO 2.OOE+OO 67 NU6II 2.OOE+OO 2.OOE+OO 2.OOE+OO 2 •OOE+OO 68 NU6I 1.66E+01 2.22E+01 1.66E+01 2 .22E+01 69 NU5IV 2.OOE+OO 2.OOE+OO 2.OOE+OO 2 .OOE+OO 70 NU51II 2.OOE+OO 2.OOE+OO 2.OOE+OO 2 •OOE+OO 71 NU5II 1.86E+01 2.50E+01 1.S6E+01 2. 50E+01 72 NU5I 2.13E+01 2.87E+01 2.13E+01 2. 87E+01 73 NU4IV 2.OOE+OO 2.OOE+OO 2. OOE+OO 2 .OOE+OO 74 NU4III 2.OOE+OO 2.OOE+OO 2.OOE+OO 2.OOE+OO 75 NU4II 2.76E+01 3.74E+01 2.76E+01 3.74E+01 76 NU41 3.42E+01 4.66E+01 3.42E+01 4 .66E+01 7 7 78 CONVECTION COEFFICIENTS 79 —Case 1— Case 2-- 80 OXIDIZER FUEL OXIDIZER FUEL 81 H6IV [W/m2K] O.OOE+OO O.OOE+OO O.OOE+OO 0..OOE+OO 82 H6III [W/ra2K] O.OOE+OO O.OOE+OO O.OOE+OO 0.•OOE+OO 83 H6II [W/m2K] O.OOE+OO O.OOE+OO O.OOE+OO 0,.OOE+OO 84 H6I [W/m2K] 1.22E+00 1.21E+00 1.22E+00 1,.21E+00 85 H5IV [W/m2K] O.OOE+OO O.OOE+OO O.OOE+OO 0 .OOE+OO 86 H5III [W/m2K] O.OOE+OO O.OOE+OO O.OOE+OO 0 .OOE+OO 87 H5II [W/m2K] 1.15E+00 1.14E+00 1.15E+00 1,.14E+00 88 H5I [W/m2KJ 1.08E+00 1.08E+00 1.08E+00 1,.08E+00 89 H4IV [W/m2K] O.OOE+OO O.OOE+OO O.OOE+OO 0,.OOE+OO 90 H4III lW/m2K] O.OOE+OO O.OOE+OO O.OOE+OO O.OOE+OO 91 H4II [W/m2K] 9.59E-01 9.62E-01 9.59E-01 9,.62E-01 S2 H4I [W/m2KJ 8.75E-01 8.81E-01 8.75E-01 8.81E-01 146 F ABCD E F G 94 THERMAL HEAT INPUT 95 —Case 1— Case 2-- 96 OXIDIZER FUEL OXIDIZER FUEL 97 Q6IV [W] 0.OOE+OO 0.OOE+OO 0. OOE+OO 0.OOE+OO 98 Q6III [W] 0.00E+00 0.OOE+OO 0.OOE+OO 0.OOE+OO 99 Q6II [W] 0.00E+00 0.OOE+OO 0.OOE+OO 0.OOE+OO 100 Q6I [W] 2.30E+01 6.11E+01 2.30E+01 6.11E+01 101 Q5IV [W) 0.00E+00 0.OOE+OO 0.OOE+OO 0.OOE+OO 102 Q5III [W] 0.00E+00 0.OOE+OO 0.OOE+OO 0.OOE+OO 103 Q5II [W] 3.07E+01 8.17E+01 3.07E+01 8.17E+01 104 Q5I [W] 4.31E+01 1 .15E+02 4.31E+01 1.15E+02 105 Q4IV [W] 0.OOE+OO 0.OOE+OO 0.OOE+OO 0.OOE+OO 106 Q4III [W] 0.00E+00 0.OOE+OO 0.OOE+OO 0.OOE+OO 107 Q4II [W] B.21E+01 2.20E+02 8.21E+01 2.20E+02 108 Q4I [W] 1.39E+02 3 .74E+02 1.39E+02 3.74E+02 109 110 REFRIGERATION MASS JUIRED 111 Case 1— Case 2-- 112 OXIDIZER FUEL OXIDIZER FUEL 113 MROX6IV [Kg] 0.OOE+OO 0.OOE+OO 0.OOE+OO 0.OOE+OO 114 MROX6II I [kg] 0.OOE+OO 0.OOE+OO 0.OOE+OO 0.OOE+OO 115 MROX6II (kg] 0.OOE+OO 0.OOE+OO 0.OOE+OO 0.OOE+OO 116 MROX6I [kg] 6.13E+00 1 .63E+01 6.13E+00 1.63E+01 117 MROX5IV [kg] 0.OOE+OO 0 .OOE+OO 0.OOE+OO 0.OOE+OO 118 MROX5II I [kg] 0.OOE+OO 0 .OOE+OO 0.OOE+OO 0.OOE+OO 119 MROX5II [kg] 8.18E+00 2.18E+01 8.18E+00 2.1BE+01 120 MROX5I [kg] 1.15E+01 3.07E+01 1.15E+01 3.07E+01 121 MROX4IV [kg] 0.OOE+OO 0.OOE+OO 0.OOE+OO 0. OOE+OO 122 MROX4II I [kg] 0.OOE+OO 0 •OOE+OO 0.OOE+OO 0.OOE+OO 123 MROX4II [kg] 2.19E+01 5.87E+01 2.19E+01 5.87E+01 124 MROX4I [kg] 3.71E+01 9. 97E+01 3.71E+01 9.97E+01 125 G A B C D E F X SHEET #7 - SPACECRAFT DETAILS 2 (Identical Cor each stage) 3 4 5 Case 1 Case 2 Case 3 Case 4 6 mu um:ISBBBB iKiaaiii •••IflBB 7 FUEL (outbound) H2 H2 CH4 H2 8 OXIDIZER (outbnd) 02 02 02 02 9 ISPP FUEL — — 10 ISPP OXID. 02 02 02 11 FUEL (return) H2 H2 CH4 CH4 12 OXIDIZER (return) 02 02 02 02 1 1 J 14 AC 0.150 0.150 0.150 0.150 15 AU [AU] 1.000 1.000 1.000 1.000 16 DF(outbnd) [kg/m3] 31.360 31.360 160.436 31.360 17 DF(return) [kg/m3] 31.360 31.360 160.436 31.360 18 DOX(outbd) [kg/m3] 436.110 436.110 436.110 436.110 19 DOX(retrn) [kg/m3] 436.110 436.110 436.110 436.110 20 EMISSF 0.040 0.040 0.040 0.040 21 EMISSOX 0.040 0.040 0.040 0.040 22 EC 0.027 0.027 0.027 0.027 23 FAF 0.500 0.500 0.500 0.500 24 FAOX 0.500 0.500 0.500 0.500 25 FFL(outbound) 0.200 0.200 0.010 0.200 26 FFL(return) 0.200 0.200 0.010 0.010 27 FOXL(outbound) 0.020 0.020 0.020 0.020 28 FOXL(return) 0.020 0.020 0.020 0.020 29 FSC(press .) 0.025 0.025 0.025 0.025 30 FSC(pump) 0.057 0.057 0.057 0.057 31 FSF 1.000 1.000 1.000 1.000 32 FSOX 1.000 1.000 1.000 1.000 33 FTC 1.000 1.000 1.000 1.000 34 FTR 0.700 0.700 0.700 0.700 35 GC 0.060 0.060 0.060 0.060 36 GEEZ 7.000 7.000 7.000 7.000 37 GZERO [m/s2] 9.810 9.810 9.810 9.810 36 ISP(outbound) [s] 445.950 445.950 356.050 445.950 39 ISP(return) js] 445.950 445.950 356.050 356.050 40 MFFAC [kg/10kgH2] 1008.000 1008.000 1008.000 1008.000 4: MOXFAC [kg/10kg02] 100.002 100.002 100.002 100.002 42 MROV [kg] 150.000 150.000 150.000 150.000 43 MRPF [kg/W] 0.060 0.060 0.060 0.060 44 MRPOX [kg/W] 0.060 0.060 0.060 0.060 45 MSAMPL [kg] 1.000 1.000 1.000 1.000 46 MSUPP [kg] 1.000 1.000 1.000 1.000 47 NC(outbound) 0.059 0.059 0.059 0.059 48 NC(return) 0.059 0.059 0.059 0.059 148 B C D E F NFT 4 4 4 4 NOXT 4 4 4 4 SI NRF 0.225 0.225 0.225 0.225 52 NROX 0.225 0.225 0.225 0.225 53 OXTC 1.000 1.000 1.000 1.000 54 OXTR 0.700 0.700 0.700 0.700 55 PF(outbound) 15.152 15.152 26.316 15.152 56 PF(return) 15.152 15.152 26.316 26.316 57 PSTOR [Pa] 3,.45E+06 3.45E+06 3.45E+06 3.. 45E+06 58 SBC [W/m2K41 5,. 67E-08 5•67E-08 5.67E-08 5,. 67E-08 59 SCE 0.026 0.026 0.026 0.026 60 SCS 0.032 0.032 0.032 0.032 61 SFE [W/m2] 1388.000 1388.000 1388.000 1388.000 62 TF(outbound) [K] 27.94 27.94 185.55 27.94 63 TF(return) [K] 20.22 20.22 123.82 123.82 64 TOX(outbound) [K] 149.58 149.58 149.58 149.58 65 TOX(return) [K] 98.64 98.64 98.64 98.64 66 THRUSTF 1.500 1.500 1.500 1.500 67 68 :TURAL DETAILS 69 CARBON - CARBON COMPOSITE 70 NSF 5 5 5 5 71 DNOZ [kg/m3] 1650 1650 1650 1650 72 WNOZ tm] 0.02 0.02 0.02 0.02 73 HALFANGL [deg] 18 18 18 18 74 75 TITANIUM ALLOY 76 FTSFPRESS 1 1 1 1 77 FTSFPUMP 3 3 3 3 78 DFT [Kg/m3] 7803 7803 7803 7803 79 SIGFT [N/m2] 6,, 20E+08 6 .20E+08 6.20E+08 6.,20E+08 80 81 TITANIUM ALLOY 82 OXTSFPRESS 1 1 1 1 83 OXTSFPUMP 3 3 3 3 84 DOXT [kg/m3] 7803 7803 7803 7803 85 SIGOXT [N/m2] 6..20E+08 6.20E+08 6.20E+08 6..20E+08 86 A B C D E F G 1 SHEET *8 - MODULAR ENGINE DETAILS (Identical for each stage) 2 3 COMPONENT MASSES (Based on Pressure-Feed System) 4 [kg] 5 Case 3 Case 4 6 •••••I 7 THRUST CHAMBER (ME) 0.12 0.12 8 STRUCTURE (MSE) 0.11 0.11 9 FEED SYSTEM (MFS) 0.11 0.11 10 NOZZLE (MN) 0.26 0.26 11 FUEL TANK (MFT) 0.25 0.25 12 OXIDIZER TANK (MOXT) 0.25 0.25 13 MMODMT 1.10 1.10 14 15 FUEL (MF) 0.59 0.59 16 OXIDIZER (MOX) 1.64 1.64 17 TOTAL PROPELLANT (MP) 2.23 2.23 18 19 MMOD 3.0885 3.0988 20 MODTHR 9.1118 9.1118 21 22 CALCULATED VALUES 3.3309 3.3303 23 CALCULATED VALUES 9.5915 9.5915 150 A C D E F l SHEET *9 Case 1 ALL EARTH TRANSPORTED LH2 AND LOX 2 MLEO [kg] 69168 3 4 LEG 6 HOHMANN to Highly eccentric Earth orbit 5 BCBBCIBO 6 TOTAL DELTA V 1500 TOTAL * OF STAGES 7 MODULAR ENGINES No AEROBRAKING Yes 8 OX. REFRIG. No FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML (kg] 2.00 0.00 0.00 0.00 12 DV [m/s] 1500 0 0 0 13 ISPUSED [s] 445.95 445.95 445.95 445.95 14 MRATIO 0.7097 1.0000 1.0000 1.0000 15 RBS [AU] 1 1 -1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 0.2147 HF 0.0004 21 ROXT (m/(kg)"1/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)*2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0072 QOUT 0.0001 24 QCOOL [ " " ] 6.9194 QCOOL 40.1797 25 MRPRE [kg^l/3] 1.8452 MRPRE 10.7146 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 9 0 0 0 30 TIMEB [sec] 86.30 0.00 0.00 0.00 31 OPT. FEED SYSTEM press. n/a n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0. 002293 0.002293 0.002293 0.002293 35 VF [m3/kg] 0. 031B88 0 .031888 0.031888 0.031888 36 OXTC 0.0551 0.0000 0.0000 0.0000 37 FTC 0.1369 0.0000 0.0000 0.0000 38 FSC 0.0250 0.0250 0.0250 0.0250 39 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.4204 0.2283 0.2283 0.2283 42 COEFTL 0.5704 0.2283 0.2283 0.2283 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.1394 0.7717 0.7717 0.7717 45 DENOM 4.3981 0.0000 0.0000 0.0000 46 VALUE 3.1686 0.0000 0.0000 0.0000 47 VAL2 0.4296 0.7717 0.7717 0.7717 48 V2ML 0.8593 0.0000 0.0000 0.0000 49 151 I A B 50 SOLVING CUBIC FOR MF 51 A1 -8.135E-01 0.000E+00 0.000E+00 O.OOOE+OO 52 A2 2.206E-01 O.OOOE+OO O.OOOE+OO 0.000E+00 53 A3 -1.994E-02 0.000E+00 O.OOOE+OO 0.000E+00 54 55 Q -6.023E-21 O.OOOE+OO 0.000E+00 O.OOOE+OO 56 R O.OOOE+OO 0.000E+00 O.OOOE+OO O.OOOE+OO 57 D -2.185E-61 O.OOOE+OO O.OOOE+OO O.OOOE+OO 58 S 0.0000 0.0000 0.0000 0.0000 59 T -0.0000 0.0000 0.0000 0.0000 60 MF [kg] 0.2712 0.0000 0.0000 0.0000 61 62 MASS SUMMARY (Hohmann to highly eccentric Earth orbit) 63 MA [kg] 0.62 0.00 0.00 0.00 64 ML [kg] 2.00 0.00 0.00 0.00 65 MG [kg] 0.65 0.40 0.40 0.40 66 MSS [kg] 0.13 0.00 0.00 0.00 67 MRF [kg] 0.00 0.00 0.00 0.00 68 MROX [kg] 0.00 0.00 0.00 0.00 69 70 ME [kg] 0.11 0.00 00 00 71 MSE [kg] 0.11 0.00 00 00 72 MFS [kg j 0.10 0.00 00 00 73 MN [kg] 0.24 0.00 00 00 74 MFT [kg] 0.57 0.00 00 00 75 MOXT [kg] 0.23 0.00 0.00 00 75 77 MF [kg] 27 0.00 0.00 00 78 MOX [kg] 52 0.00 0.00 00 79 MP [kg] 79 0.00 0.00 00 80 81 NUMMOD 82 MMODTL [kg] 83 MODTHRTL [kg] 84 85 Ml [kg] 4.78 0.40 0.40 0.40 86 MO [kg] 6.17 0.00 0.00 0.00 87 152 I H I J K L M 1 Case 1 2 3 4 LEG 5 MPO to HOHMANN towards Earth 5 CCOCBinaBCCBCnBICBaaBCBBBrilB.irBnBEEBBrilBBUBaBBllBBBOBITBnnBnBBaEC 6 TOTAL DELTA V 2700 TOTAL • OF STAGES 2 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. No FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 11.41 6.25 0.00 0.00 12 DV [m/s] 1350 1350 0 0 13 ISPUSED [s] 445.95 445.95 445.95 445.95 14 MRAT10 0.7345 0.7345 1.0000 1.0000 15 RBS [AU) 1111 16 SFB [W/m2] 1388 1388 1388 13B8 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 0.2147 HF 0.0004 21 ROXT [m/(kg)-l/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)*2/3] 1.7370 QIN 10.0450 23 QOUT [ " " j 0.0072 QOUT 0.0001 24 QCOOL [ " " ] 6.9194 QCOOL 40.1797 25 MRPRE [kg~l/3] 1.8452 MRPRE 10.7146 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 31 17 0 0 30 TIMEB [sec] 78.94 78.94 0.00 0.00 31 OPT. FEED SYSTEM press. press . n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0.002293 0.002293 0,.002293 0..002293 35 VF [m3/kg] 0.031888 0.031888 0..031888 0..031888 36 OXTC 0.0504 0.0504 0.0000 0.0000 37 FTC 0.1252 0.1252 0.0000 0.0000 38 FSC 0.025 0.025 0.025 0.025 39 4 0 INTERMEDIATE MASS CALCS 41 COEFSUM 0.4040 0.4040 0.2283 0.2283 42 COEFTL 0.4040 0.4040 0.2283 0.2283 43 COEFTLM 0.0000 0.0000 0..0000 0.0000 44 NUMER 0.3305 0.3305 0.7717 0.7717 45 DENOM 4.0230 4.0230 0.0000 0.0000 46 VALUE 8.2151 8.2151 0.0000 0.0000 47 VAL2 0.5960 0.5960 0.7717 0.7717 48 V2ML 6.8027 3.7254 0.0000 0.0000 49 153 I H I J K L M 50 SOLVING CUBIC FOR MF 51 A1 -2.484E+00 -1,. 360E+00 0,•OOOE+OO 0. OOOE+OO 52 A2 2.057E+00 6 .169E-01 0,.OOOE+OO O.OOOE+OO 53 A3 -5.678E-01 -9,.326E-02 0..OOOE+OO 0.OOOE+OO 54 55 Q -9.637E-20 0 .OOOE+OO 0,•OOOE+OO 0.OOOE+OO 56 R 9.637E-20 0..OOOE+OO 0..OOOE+OO O.OOOE+OO 57 D 9.288E-39 0,.OOOE+OO 0.•OOOE+OO O.OOOE+OO 58 S 0.0000 0.0000 0.0000 0.0000 59 T 0.0000 0.0000 0.0000 0.0000 60 MF [kg] 0.8281 0.4535 0.0000 0.0000 £D 11 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 00 64 ML [kg] 11.41 6.25 0.00 00 65 MG [kg] 0.95 0.70 0.40 40 66 MSS [kg] 0.29 0.16 0.00 00 67 MRF [kg] 0.00 0.00 0.00 00 68 MROX [kg] 0.00 0.00 0.00 00 69 70 ME [kg] 0.24 0.13 0.00 00 71 MSE [kg] 0.24 0.13 0.00 00 72 MFS [kg] 0.23 0.13 0.00 00 73 MN [kg] 0.54 0.30 0.00 00 74 MFT [kg] 1.15 0.63 0.00 00 75 MOXT [kg] 0.46 0.25 0.00 00 76 77 MF [kg] 0.83 0.45 0.00 0.00 78 MOX [kg] 4.64 2.54 0.00 0.00 79 MP [kg] 5.47 2.99 0.00 0.00 80 81 NUMMOD 82 MMODTL [kg] 83 MODTHRTL [kg] 84 85 Ml [kg] 15.52 8.68 0.40 0.40 86 MO [kg] 20.58 11.27 0.00 0.00 87 154 I Q 1 Case 1 2 3 4 LEG 4 Mars Ascent Vehicle to MPO 5 6 TOTAL DELTA V 3500 TOTAL # OF STAGES 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. NO FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg) 52.34 20.84 0.00 0.00 12 DV [m/s] 1750 1750 0 0 13 ISPUSED [s] 445.95 445.95 445.95 445.95 14 MRATIO 0.6703 0.6703 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX (W/m2] 0.2147 HF 0.0004 21 ROXT [m/(kg)~l/3] 0.0515 RFT 0.1239 22 QIN [W/(kgr2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0072 QOUT 0.0001 24 QCOOL [ " " j 6.9194 QCOOL 40.1797 25 MRPRE [kg^l/3] 1.8452 MRPRE 10.7146 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 194 77 0 0 30 TIMEB [secj . 98.02 98.02 0.00 0.00 31 OPT. FEED SYSTEM press. press. n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0.002293 0 .002293 0.002293 0..002293 35 VF [m3/kg] 0.031888 0 .031888 0.031888 0..031888 36 OXTC 0.0626 0.0626 0.0000 0.0000 37 FTC 0.1555 0.1555 0.0000 0.0000 38 FSC 0.025 0.025 0.025 0.025 39 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.4465 0.4465 0..2283 0.2283 42 COEFTL 0.4465 0.4465 0,.2283 0.2283 43 COEFTLM 0.0000 0.0000 0,.0000 0.0000 44 NUMER 0.2239 0.2239 0,.7717 0.7717 45 DENOM 4.9954 4.9954 0..0000 0.0000 46 VALUE 4.4812 4.4812 0 .0000 0.0000 47 VAL2 0.5535 0.5535 0 .7717 0.7717 48 V2ML 28.97 11.54 0.00 0.00 49 155 i 0 P 50 SOLVING CUBIC FOR MF 51 A1 -1.940E+01 -7.724E+00 0.OOOE+OO 0. OOOE+OO 52 A2 1.254E+02 1.988E+01 0.OOOE+OO 0.OOOE+OO 53 A3 -2.703E+02 -1.706E+01 0.000E+00 0.OOOE+OO 54 55 Q 1.234E-17 -3.855E-19 0.000E+00 0.OOOE+OO 56 R 8.224E-17 -4.112E-18 0.000E+00 0.OOOE+OO 57 D 6.763E-33 1.691E-35 0.000E+00 0.OOOE+OO 58 S 0.0000 0.0000 0.0000 0.0000 59 T 0.0000 -0.0000 0.0000 0.0000 60 MF [kg] 6.4655 2.5745 0.0000 0.0000 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kg] 52.34 20.84 0.00 0.00 65 MG [kg] 5.03 2.24 0.40 40 66 MSS [kg] 2.47 0.98 0.00 00 67 MRF [kg] 0.00 0.00 0.00 00 68 MROX [kg] 0.00 0.00 0.00 00 69 70 ME [kg] 05 81 0.00 00 71 MSE [kg] 00 80 0.00 00 72 MFS [kg] 93 77 0.00 00 73 MN [kg] 53 80 0.00 00 74 MFT [kg] 11.99 77 0.00 00 75 MOXT [kg] 4.83 92 0.00 00 76 77 MF [kg] 6.47 2.57 00 00 78 MOX [kg] 36.21 14.42 00 00 79 MP [kg] 42 .67 16.99 00 0.00 80 81 NUMMOD 82 MMODTL [kg] 83 MODTHRTL [kg] 84 85 Ml [kg] 87.16 34.95 0.40 0.40 86 MO [kg] 129.43 51.54 0.00 0.00 87 156 X Y AA l Case 1 2 LEG 3 MFO to Mars Soft Landing 3 BBSSBSBIBBB83IB8BB8B1 ••BBSBKI 4 TOTAL DELTA V 1000 TOTAL # OF STAGES 5 NODULAR ENGINES NO AEROBRAKING Yes 6 OX. REFRIG. Yes FUEL REFRIG. Yes 7 OX. SURF. REF. Yes FUEL SURF. REF. Yes 8 OX. ISRU No FUEL ISRU No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 593.43 0.00 0.00 0.00 12 DV [m/s] 1000 0 0 0 13 ISPUSED [s] 445.95 445.95 445.95 445.95 14 MRATIO 0.7957 1.0000 1.0000 1.0000 15 RBS [AU) 1 1 -1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 1.1352 HF 0.0014 21 ROXT [m/(kg)"l/3] 0.0515 RFT 0.1239 22 QIN (W/(kg)*2/3] 1.7370 QIN 10.0450 23 QOUT [ » " ] 0.0379 QOUT 0.0003 24 QCOOL [ " " ] 6.7965 QCOOL 40.1789 25 MRPRE [kg-1/3] 1.8124 MRPRE 10.7144 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 2365 0 0 0 30 TIMEB [secj 60.75 0.00 0.00 0.00 31 OPT. FEED SYSTEM pump n/a n/a n/a 32 PT [MPa] 0.3447 0.3447 0.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0.000934 0,.000934 0.000934 0 .000934 35 VF [m3/kg] 0.015592 0,.015592 0.015592 0,.015592 36 OXTC 0.0032 0.0000 0.0000 0.0000 37 FTC 0.0094 0.0000 0.0000 0.0000 38 FSC 0.0833 0.0833 0.0833 0.0833 J 7 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.2992 0.2866 0.2866 0.2866 42 COEFTL 0.4492 0.2866 0.2866 0.2866 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3464 0.7134 0.7134 0.7134 45 DENOM 3.0961 0.0000 0.0000 0.0000 46 VALUE 11.1895 0.0000 0.0000 0.0000 47 VAL2 0.5508 0.7134 0.7134 0.7134 48 V2ML 326.84 0.00 0.00 0.00 49 157 I V W X Y Z AA 50 SOLVING CUBIC FOR MF 51 A1 -9.080E+01 0.OOOE+OO 0.OOOE+OO O.OOOE+OO 52 A2 2.560E+03 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 53 A3 -2.492E+04 0.OOOE+OO O.OOOE+OO O.OOOE+OO 54 55 Q -6.276E+01 O.OOOE+OO O.OOOE+OO O.OOOE+OO 56 R 1.449E+03 O.OOOE+OO O.OOOE+OO O.OOOE+OO 57 D 1.853E+06 O.OOOE+OO O.OOOE+OO O.OOOE+OO 58 S 14.1123 0.0000 0.0000 0.0000 59 T 4.4471 0.0000 0.0000 0.0000 60 MF [kg] 48.8245 0.0000 0.0000 0.0000 61 62 MASS SUMMARY 63 MA [kg] 147.53 0.00 0.00 0.00 64 ML [kg] 593.43 0.00 0.00 0.00 65 MG [kg] 59.41 0.40 0.40 0.40 66 MSS [kg] 31.47 0.00 0.00 0.00 67 MRF [kgj 163.15 0.00 0.00 0.00 68 MROX [kg] 87.03 0.00 0.00 0.00 69 70 ME [kg] 26.11 0.00 0.00 0.00 71 MSE [kg] 25.57 0.00 0.00 0.00 72 MFS [kgj 81.94 0.00 0.00 0.00 73 MN [kg] 57.82 0.00 0.00 0.00 74 MFT [kg] 9.27 0.00 0.00 0.00 75 MOXT [kg] 3.11 0.00 0.00 0.00 76 77 MF [kg] 48.82 0.00 0.00 0.00 78 MOX [kg] 273.42 0.00 0.00 0.00 79 MP [kg] 322.24 0.00 0.00 0.00 80 81 82 83 84 85 XI [kg] 1285.84 0.40 0.40 0.40 86 MO [kg] 1576.99 0.00 0.00 0.00 87 158 AC AD AE AF AG AH 1 Case 1 2 3 4 LEG 2 HOHMANN to Mars Parking Orbit (200 km) 5 6 TOTAL DELTA V 2700 TOTAL * OF STAGES 2 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. Yes FUEL REFRIG. Yes 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 3672.95 1592.22 0.00 0.00 12 DV [m/s] 1850 1350 0 0 13 ISPUSED [S] 445.95 445.95 445.95 445.95 14 MRATIO 0.6552 0.7345 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 1.1352 HF 0.0014 21 ROXT [m/(kg)"1/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)~2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0379 QOUT 0.0003 24 QCOOL ( " " ] 6.7 965 QCOOL 40.1789 25 MRPRE [kg~l/3] 1.8124 MRPRE 10.7144 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 15733 5441 0 0 30 TIMEB [sec] 102.52 78.94 0.00 0.00 31 OPT. FEED SYSTEM pump pump n/a n/a 32 PT [MPa] 0.3447 0.3447 0.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0.000934 0.000934 0.000934 0.000934 35 VF [m3/kg] 0.015592 0.015592 0.015592 0.015592 36 OXTC 0.0053 0.0041 0.0000 0.0000 37 FTC 0.0159 0.0122 0.0000 0.0000 38 FSC 0.0833 0.0833 0.0833 0.0833 39 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3079 0.3030 0.2866 0.2866 42 COEFTL 0.3079 0.3030 0.2866 0.2866 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3473 0.4315 0.7134 0.7134 45 DENOM 5.2249 4.0230 0.0000 0.0000 46 VALUE 6.6466 10.7256 0.0000 0.0000 47 VAL2 0.6921 0.6970 0.7134 0.7134 48 V2ML 2542.11 1109.79 0.00 0.00 49 159 i AC AO AE AF AG AH 50 SOLVING CUBIC FOR MF 51 A1 -1.163E+03 -3.140E+02 0 • OOOE+OO 0.OOOE+OO 52 A2 4.388E+05 3.212E+04 0. OOOE+OO 0.OOOE+OO 53 A3 -5.595E+07 •1.108E+06 0.000E+00 0.OOOE+OO 54 55 Q -3.877E+03 •2.494E+02 0.OOOE+OO 0.OOOE+OO 56 R 1.134E+06 1.969E+04 0.OOOE+OO 0.OOOE+OO 57 D 1.228E+12 3.721E+08 0.OOOE+OO 0.OOOE+OO 58 S 130.8803 33.9061 0.0000 0.0000 59 T 29.6196 7.3550 0.0000 0.0000 60 MF [kg] 548.00 145.93 0.00 0.00 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kg] 3672.95 1592.22 0.00 0.00 65 MG [kg] 409.32 122.51 0.40 0.40 66 MSS [kg] 218.09 65.13 0.00 0.00 67 MRF [kg] 726.72 311.18 0.00 0.00 68 MROX [kg] 387.65 166.00 0.00 0.00 69 70 ME [kg] 180.95 54.03 00 0.00 71 MSE [kg] 177.20 52.92 00 0.00 72 MFS [kg] 567.78 169.55 00 0.00 73 MN [kg] 400.62 119.63 00 0.00 74 MFT [kg] 108.40 24.92 00 0.00 75 MOXT [kg] 36.36 8.36 00 0.00 76 77 MF [kg] 548.00 145.93 00 0.00 78 MOX [kg] 3068.82 817.21 00 0.00 79 MP [kg] 3616.83 963.14 00 0.00 80 81 82 83 84 85 Ml [kg] 6886.05 2686.46 0.40 0.40 86 MO [kg] 10488.35 3627.42 0.00 0.00 87 160 AJ AK AL AM AM AO l Case 1 2 3 4 LEG 1 LEO to HOHMANN towards Mars 5 6 TOTAL DELTA V 3800 TOTAL * OF STAGES 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. Yes FUEL REFRIG. Yes 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 28529.43 10659.33 0.00 0.00 12 DV [m/s] 1900 1900 0 0 13 ISPUSED [s] 445.95 445.95 445.95 445.95 14 MRATIO 0.6477 0.6477 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB (W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 1.1352 HF 0.0014 21 ROXT [m/(kg)*l/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)~2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0379 QOUT 0.0003 24 QCOOL [ " " ] 6.7965 QCOOL 40.1789 25 MRPRE [kg~l/3] 1.8124 MRPRE 10.7144 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 104665 42093 0 0 30 TIMEB [sec] 104.74 104.74 0.00 0.00 31 OPT. FEED SYSTEM pump pump n/a n/a 32 PT [MPa] 0.3447 0.3447 0.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0,.000934 0.000934 0..000934 0..000934 35 VF [m3/kg] 0,.015592 0.015592 0..015592 0 .015592 36 OXTC 0.0055 0.0055 0.0000 0.0000 37 FTC 0.0162 0.0162 0.0000 0.0000 38 FSC 0.0833 0.0833 0.0833 0.0833 J1Q 7 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3083 0.3083 0.2866 0.2866 42 COEFTL 0.3083 0.3083 0.2866 0.2866 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3394 0.3394 0.7134 0.7134 45 DENOM 5.3377 5.3377 0.0000 0.0000 46 VALUE 6.3580 6.3580 0.0000 0.0000 47 VAL2 0.6917 0.6917 0.7134 0.7134 48 V2ML 19732.66 7372.63 0.00 0.00 49 161 I AJ AK AL AM AN AO 50 SOLVING CUBIC 51 A1 -9.328E+03 -3.496E+03 0.OOOE+OO 0.OOOE+OO 52 A2 2.B90E+07 4.034E+06 0.OOOE+OO 0.OOOE+OO 53 A3 -2.9B9E+10 -1.559E+09 0.000E+00 0.OOOE+OO 54 55 Q -3.574E+04 -1.337E+04 0.OOOE+OO 0.OOOE+OO 56 R 8.341E+07 1.172E+07 0.OOOE+OO 0.OOOE+OO 57 D 6.912E+15 1.349E+14 0.OOOE+OO 0.OOOE+OO 58 S 550.1929 285.7404 0.0000 0.0000 59 T 64.9508 46.7991 0.0000 0.0000 60 MF [kg] 3724.48 1497.87 0.00 0.00 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kgj 28529.43 10659.33 0.00 0.00 65 MG [kg] 2475.25 1044.57 0.40 0.40 66 MSS [kg] 1319.92 556.89 0.00 0.00 67 MRF [kg] 2579.28 1409.26 0.00 0.00 68 MROX [kg] 1375.87 751.74 0.00 0.00 69 70 ME [kg] 1095.12 462.04 0.00 0.00 71 MSE [kg] 1072.43 452.47 0.00 0.00 72 MFS [kg] 3436.27 1449.80 0.00 0.00 73 MN [kg] 2424.60 1022.97 0.00 0.00 74 MFT [kg] 670.23 282.78 0.00 0.00 75 MOXT [kg] 224.83 94.86 0.00 0.00 76 77 MF [kg] 3724.4B 1497.87 0.00 0.00 78 MOX [kg] 20857.06 8388.06 0.00 0.00 79 MP [kg] 24581.54 9885.93 0.00 0.00 80 81 82 83 84 85 Ml [kg] 45203.23 18186.69 0.40 0.40 86 MO [kg] 69776.89 28062.09 0.00 0.00 87 162 I A B C D E F 68 Case 1 89 90 91 MARS SURFACE DETAILS 92 93 MARS ROVER [kg] 150 94 OX. FACTORY [kg] 100.00 > MASS USED 0.00 95 FUEL FACTORY [kg] 1008.00 > MASS USED 0.00 70OA 97 OXIDIZER REFRIGERATION UNITS [kg] 98 STAGE I STAGE II STAGE III STAGE IV 99 LEG 6 6.1334 0.0000 0.0000 0.0000 100 LEG 5 11.4899 8.1817 0.0000 0.0000 101 LEG 4 37.1340 21.9007 0.0000 . 0.0000 102 103 TOTAL 84.8397 > MASS USED 84.8397 104 105 FUEL REFRIGERATION UNITS [kg] 106 STAGE I STAGE II STAGE III STAGE IV 107 LEG 6 16.2838 0.0000 0.0000 0.0000 108 LEG 5 30.6726 21.7818 0.0000 0.0000 109 LEG 4 99.7123 58.6890 0.0000 0.0000 110 111 TOTAL 227.1394 > MASS USED 227.1394 112 113 OXIDIZER NECESSARY FOR RETURN [kg] 114 STAGE I STAGE II STAGE III STAGE IV 115 LEG 6 1.5186 0.0000 0.0000 0.0000 116 LEG 5 4.6372 2.5395 0.0000 0.0000 117 LEG 4 36.2066 14 .4172 0.0000 0.0000 118 119 TOTAL 59.3192 > MASS USED 0.0000 120 121 FUEL NECESSARY FOR RETURN [kg] 122 STAGE I STAGE II STAGE III STAGE IV 123 LEG 6 0.2712 0.0000 0.0000 0.0000 124 LEG 5 0.8281 0.4535 0.0000 0.0000 125 LEG 4 6.4655 2.5745 0.0000 0.0000 126 127 TOTAL 10.5927 > MASS USED 0.0000 128 163 I HI J K L M 89 TOTAL MASS SUMMARY Case 1 90 91 PRE-MARS POST-MARS 92 MA [kg] 148 MA [kg] 0.6 93 MG [kg] 4114 MG [kg] 12.4 94 MSS [kg] 2192 MSS [kg] 4.0 95 MRF [kg] 5190 MRF [kg] 0.0 96 MROX [kg] 2768 MROX [kg] 0.0 97 98 ME [kg] 1818 ME [kg] 3.3 99 MSE [kg] 1781 MSE [kg] 3.3 100 MFS [kg] 5705 MFS [kg] 3.2 101 MN [kg] 4026 MN [kg] 7.4 102 MFT [kg] 1096 MFT [kg] 19.1 103 MOXT [kg] 368 MOXT [kg] 7.7 104 105 MF [kg] 5965 MF [kg] 10.6 106 MFL [kg] 448.1 MFL [kg] 2.1 107 MOX [kg] 33405 MOX [kg] 59.3 10B MOXL [kg] 251.0 MOXL [kg] 1.2 109 MP [kg] 39370 MP [kg] 69.9 110 111 ML [kg] 29204 ML [kg] 61.0 112 I 0 P Q R S T 68 Case 1 89 90 TOTAL MISSION 91 MA (kg] 148.16 ME [kg] 1821.61 92 MG [kg] 4126.23 MSE [kg] 1783.87 93 MSS [kg] 2195.54 MFS [kg] 5708.49 94 MRF [kg] 5416.73 MN [kg] 4033.06 95 MROX [kg] 2853.13 MFT [kg] 1114.71 96 MOXT [kg] 375.22 97 MF [kg] 5975.69 98 MFL [kg] 450.24 99 MOX [kg] 33463.89 100 MOXL [kg] 252.14 MSAMPL [kg] 1.00 101 MP [kg] 39439.58 MSUPP [kg] 1.00 102 ML [kg] 29576.74 MROV [kg] 150.00 103 MOXFAC [kg] 0.00 104 MFFAC [kg] 0.00 105 106 107 MLEO [kg] 69168.32 108 165 JAB C D E F 1 SHEET #10 Case 2 ALL EARTH TRANSPORTED LH2 AND ISR 2 MLEO [kg] 66356 3 4 LEG 6 HOHMANN to Highly eccentric Earth orbit 6 TOTAL DELTA V 1500 TOTAL # OF STAGES 1 7 MODULAR ENGINES No AEROBRAKING Yes 8 OX. REFRIG. No FUEL REFRIG. No 9 10 STAGE X STAGE II STAGE III STAGE IV 11 ML [kg] 2.00 0.00 0.00 0.00 12 DV [m/s] 1500 0 0 0 13 ISPUSED [s] 445.95 445.95 445.95 445.95 14 MRATIO 0.7097 1.0000 1.0000 1.0000 15 RBS [AU] 1111 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 0.2147 HF 0.0004 21 ROXT [tn/(kg)"l/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)~2/3] 1.7370 QIN 10.0450 2 3 QOUT [ " " ) 0.0072 QOUT 0.0001 24 QCOOL [ " " ] 6.9194 QCOOL 40.1797 25 MRPRE [kg^l/3] 1.8452 MRPRE 10.7146 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 9 0 0 0 30 TIMEB [sec] 86 .30 0.00 0.00 0.00 31 OPT. FEED SYSTEM press. n/a n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0.002293 0.002293 0,.002293 0.002293 35 VF [m3/kg] 0 .031888 0.031888 0,.031888 0.031888 36 OXTC 0.0551 0.0000 0.0000 0.0000 37 FTC 0.1369 0.0000 0.0000 0.0000 38 FSC 0.0250 0.0250 0.0250 0.0250 J1 7Q 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.4204 0.2203 0.2283 0.2283 42 COEFTL 0.5704 0.2283 0.2283 0.2283 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.1394 0.7717 0.7717 0.7717 45 DENOM 4.3981 0.0000 0.0000 0.0000 46 VALUE 3.1686 0.0000 0.0000 0.0000 47 VAL2 0.4296 0.7717 0.7717 0.7717 48 V2ML 0.8593 0.0000 0.0000 0.0000 49 166 J A B C 50 SOLVING CUBIC FOR MF 51 A1 -8.135E-01 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 52 A2 2.206E-01 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 53 A3 -1.994E-02 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 54 55 Q -6.023E-2I 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 56 R 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 57 D -2•185E-61 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 58 S 0.0000 0.0000 0.0000 0.0000 59 T -0.0000 0.0000 0.0000 0.0000 60 MF [kg] 0.2712 0.0000 0.0000 0.0000 61 62 MASS SUMMARY (Hohnann tcto highly eccentric Earth orbit) 63 MA [kg] 0.62 00 0.00 0.00 64 ML [kg] 2.00 00 0.00 0.00 65 MG [kgj 0.65 40 0.40 0.40 66 MSS [kg] 0.13 00 0.00 0.00 67 MRF [kgj 0.00 00 0.00 0.00 68 MROX [kg] 0.00 00 0.00 0.00 69 70 ME [kg] 0.11 0.00 00 0.00 71 MSE [kg] 0.11 0.00 00 0.00 72 MFS [kg] 0.10 0.00 00 0.00 73 MN [kg] 0.24 0.00 00 0.00 74 MFT [kg] 0.57 0.00 00 0.00 75 MOXT [kg] 0.23 0.00 00 0.00 76 77 MF [kg] 0.27 00 00 00 78 MOX [kg] 1.52 00 00 00 79 MP [kg] 1.79 00 00 00 80 81 NUMMOD 62 MMODTL [kg] 83 KODTHRTL [kg] 84 85 Ml [kg] 4 .78 0.40 0.40 0.40 86 MO [kg] 6.17 0.00 0.00 0.00 167 J l Case 2 2 3 4 LEG 5 MPO to HOHMANN towards Garth 5 6 TOTAL DELTA V 2700 TOTAL * OF STAGES 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. No FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 11.41 6.25 0.00 0.00 12 DV [m/s] 1350 1350 0 0 13 ISPUSED [s] 445.95 445.95 445.95 445.95 14 MRATIO 0.7345 0.7345 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1386 1388 1388 1388 17 IB 'RIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 0.2147 HF 0.0004 21 ROXT [m/(kg)~l/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)*2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0072 QOUT 0.0001 24 QCOOL [ " " ] 6.9194 QCOOL 40.1797 25 MRPRE [kg~l/3] 1.8452 MRPRE 10.7146 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 31 17 0 0 30 TIMEB [sec] 78.94 78.94 0.00 0.00 31 OPT. FEED SYSTEM press. press. n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0 .002293 0 .002293 0.002293 0.002293 35 VF [m3/kg] 0,.031888 0.031888 0.031888 0.031888 36 OXTC 0.0504 0.0504 0.0000 0.0000 37 FTC 0.1252 0.1252 0.0000 0.0000 38 FSC 0.0250 0.0250 0.0250 0.0250 1J 7Q 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.4040 0.4040 0.2283 0.2283 42 COEFTL 0.4040 0.4040 0.2283 0.2283 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3305 0.3305 0.7717 0.7717 45 DENOM 4.0230 4.0230 0.0000 0.0000 46 VALUE 8.2151 8.2151 0.0000 0.0000 47 VAL2 0.5960 0.5960 0.7717 0.7717 48 V2ML 6.8027 3.7254 0.0000 0.0000 49 168 J H I J K L M 50 SOLVING CUBIC FOR MF 51 A1 -2.484E+00 -X.360E+00 O.OOOE+OO O.OOOE+OO 52 A2 2.057E+00 6.169E-01 O.OOOE+OO O.OOOE+OO 53 A3 -5.678E-01 -9.326E-02 O.OOOE+OO O.OOOE+OO 54 55 Q -9.637E-20 O.OOOE+OO O.OOOE+OO O.OOOE+OO 56 R 9.637E-20 O.OOOE+OO O.OOOE+OO O.OOOE+OO 57 D 9.288E-39 O.OOOE+OO O.OOOE+OO O.OOOE+OO 58 S 0.0000 0.0000 0.0000 0.0000 59 T 0.0000 0.0000 0.0000 0.0000 60 MF [kg] 0.8281 0.4535 0.0000 0.0000 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kg] 11.41 6.25 0.00 0.00 65 MG [kg] 0.95 0.70 0.40 0.40 66 MSS [kg] 0.29 0.16 0.00 0.00 67 MRF [kg] 0.00 0.00 0.00 0.00 68 MROX [kg] 0.00 0.00 0.00 0.00 69 70 ME [kg] 0.24 0.13 0.00 0.00 71 MSE [kg] 0.24 0.13 0.00 0.00 72 MFS [kg] 0.23 0.13 0.00 0.00 73 MN [kg] 0.54 0.30 0.00 0.00 74 MFT [kg] 1.15 0.63 0.00 0.00 75 MOXT [kg] 0.46 0.25 0.00 0.00 76 77 MF [kg] 0.83 0.45 0.00 0.00 78 MOX [kg] 4.64 2.54 0.00 0.00 79 MP [kg] 5.47 2.99 0.00 0.00 80 81 NUMMOD 82 MMODTL [kg] 83 MODTHRTL [kg] 84 85 Ml [kg] 15.52 8.68 0.40 0.40 86 MO [kg] 20.58 11.27 0.00 0.00 87 169 Q l Case 2 2 3 4 LEG 4 Mars Ascent Vehicle to MPO 5 • BVBBB HMIIB 6 TOTAL DELTA V 3500 TOTAL # OF STAGES 7 MODULAR ENGINES NO AEROBRAKING No e OX. REFRIG. No FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 52.34 20.84 0.00 0.00 12 DV [m/s] 1750 1750 0 0 13 ISPUSED [sj 445.95 445.95 445.95 445.95 14 MRATIO 0.6703 0.6703 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 0.2147 HF 0.0004 21 ROXT [m/(kg)~l/3] 0.0515 RFT 0.1239 22 QIN |W/(kg)^2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0072 QOUT 0.0001 24 QCOOL [ " " ] 6.9194 QCOOL 40.1797 25 MRPRE [kg~l/3] 1.8452 MRPRE 10.7146 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 194 77 0 0 30 TIMEB [secj 98.02 98.02 0.00 0.00 31 OPT. FEED SYSTEM press. press. n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0,.002293 0.002293 0,.002293 0 .002293 35 VF [m3/kg] 0,.031888 0.031888 0,.031888 0,.031888 36 OXTC 0.0626 0.0626 0.0000 0.0000 37 FTC 0.1555 0.1555 0.0000 0.0000 38 FSC 0.0250 0.0250 0.0250 0.0250 J 7 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.4465 0.4465 0.2283 0.2283 42 COEFTL 0.4465 0.4465 0.2283 0.2263 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.2239 0.2239 0.7717 0.7717 45 DENOM 4.9954 4.9954 0.0000 0.0000 46 VALUE 4.4812 4.4812 0.0000 0.0000 47 VAL2 0.5535 0.5535 0.7717 0.7717 48 V2ML 28.97 11.54 0.00 0.00 49 170 o P SOLVING CUBIC FOR MF 51 A1 -1.940E+01 724E+00 0.000E+00 0.000E+00 52 A2 1.254E+02 988E+01 O.OOOE+OO 0.000E+00 53 A3 -2.703E+02 706E+01 O.OOOE+OO O.OOOE+OO 54 55 Q 1.234E-17 -3.855E-19 O.OOOE+OO O.OOOE+OO 56 R 8.224E-17 -4.112E-18 O.OOOE+OO O.OOOE+OO 57 D 6.763E-33 1.691E-35 O.OOOE+OO 0.000E+00 58 S 0.0000 0.0000 0.0000 0.0000 59 T 0.0000 -0.0000 0.0000 0.0000 60 MF [kg] 6.4655 2.5745 0.0000 0.0000 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kgJ 52.34 20.84 0.00 0.00 65 MG [kg] 03 2.24 0.40 0.40 66 MSS [kg] 47 0.98 0.00 0.00 67 MRF [kg] 00 0.00 0.00 0.00 68 MROX [kg] 00 0.00 0.00 0.00 69 70 ME [kg] 05 0.81 00 0.00 71 MSE [kg] 00 0.80 00 0.00 72 MFS [kgj 93 0.77 00 0.00 73 MN [kg] 53 1.80 00 0.00 74 MFT [kg] 11.99 4.77 00 0.00 75 MOXT [kg] 4.83 1.92 00 0.00 76 77 MF [kg] 6.47 2.57 0.00 0.00 78 MOX [kg] 36.21 14.42 0.00 0.00 79 MP (kg) 42.67 16.99 0.00 0.00 80 81 NUMMOD 82 MKODTL [kg] 83 MODTHRTL [kg] 84 85 Ml [kg] 87.16 34.95 0.40 0.40 86 MO [kgj 129.43 51.54 0.00 0.00 87 171 x y AA l Cose 2 2 LEG 3 MPO to Mars Soft Landing 3 4 TOTAL DELTA V 1000 TOTAL « OF STAGES 5 MODULAR ENGINES No AEROBRAKING Yes 6 OX. REFRIG. Yes FUEL REFRIG. Yes 7 OX. SURF. REF. No FUEL SURF. REF. Yes 8 OX. ISRU Yes FUEL ISRU No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 549.27 0.00 0.00 0.00 12 DV [m/s] 1000 0 0 0 13 ISPUSED [s] 445.95 445.95 445.95 445.95 14 MRATIO 0.7957 1.0000 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 1.1352 HF 0. 0014 21 ROXT [m/(kgPl/3] 0.0515 RFT 0, 1239 22 QIN [W/(kg)~2/3] 1.7370 QIN 10. 0450 23 QOUT [ " " j 0.0379 QOUT 0. 0003 24 QCOOL [ " " ] 6.7965 QCOOL 40. 1789 25 MRPRE [kg~l/3] 1.8124 MRPRE 10. 7144 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 2221 0 0 0 30 TIMEB [sec] 60.75 0.00 0.00 0.00 31 OPT. FEED SYSTEM pump n/a n/a n/a 32 PT [MPa] 0.3447 0.3447 0.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0.000934 0..000934 0 .000934 0,.000934 35 VF [m3/kg] 0.015592 0,.015592 0 .015592 0,.015592 36 OXTC 0.0032 0.0000 0.0000 0.0000 37 FTC 0.0094 0.0000 0.0000 0.0000 38 FSC 0.0833 0.0833 0.0833 0.0833 J y 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.2992 0.2866 0.2866 0.2866 42 COEFTL 0.4492 0.2866 0.2866 0.2866 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3464 0.7134 0.7134 0.7134 45 DENOM 3.0961 0.0000 0.0000 0.0000 46 VALUE 11.1895 0.0000 0.0000 0.0000 47 VAL2 0.5508 0.7134 0.7134 0.7134 48 V2ML 302.52 0.00 0.00 0.00 49 172 v w AA SOLVING CUBIC FOR MF A1 -8.427E+01 0.000E+00 0.000E+00 0.OOOE+OO A2 2•193E+03 0.000E+00 0.OOOE+OO 0. OOOE+OO A3 -1.976E+04 0.OOOE+OO 0.OOOE+OO 0. OOOE+OO 54 55 0 -5.817E+01 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 56 R 1.248E+03 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 57 0 1.362E+06 0.OOOE+OO 0.OOOE+OO 0.OOOE+OO 58 S 13.4174 0.0000 0.0000 0.0000 59 T 4.3356 0.0000 0.0000 0.0000 60 MF [kg] 45.8446 0.0000 0.0000 0.0000 61 62 MASS SUMMARY 63 MA [kg] 139.72 0.00 ,00 0.00 64 ML [kg] 549.27 0.00 . 00 0.00 65 MG [kg] 56.29 0.40 , 40 0.40 66 MSS [kg] 29.81 0.00 ,00 0.00 67 MRF [kg] 157.64 0.00 00 0.00 68 MROX [kg] 73.21 0.00 ,00 0.00 69 70 ME [kg] 24.73 00 ,00 ,00 71 MSE [kg] 24.22 00 ,00 ,00 72 MFS [kg] 77.60 00 ,00 ,00 73 MN [kg] 54 .75 00 ,00 ,00 74 MFT [kg] 8.78 00 ,00 ,00 75 MOXT [kg] 2.94 00 00 .00 76 77 MF [kg] 45.84 00 0.00 ,00 78 MOX [kg] 256.73 00 0.00 ,00 79 MP [kg] 302.57 00 0.00 0.00 80 81 82 83 84 85 Ml [kg] 1198.96 0.40 0.40 0.40 86 MO [kg] 1480.74 0.00 0.00 0.00 87 173 J AC AO AE AF AG AH 1 Case 2 2 3 4 LEG 2 HOHMANN to Mars Parking Orbit (200 km) 0e 6 TOTAL DELTA V 2700 TOTAL # OF !STAGES 2 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. Yes FUEL REFRIG Yes a 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 3475.31 1495.04 0.00 0.00 12 DV [m/s] 1850 1350 0 0 13 ISPUSED (s] 445.95 445.95 445.95 445.95 14 MRATIO 0.6552 0.7345 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 1 7/ 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 1.1352 HF 0.0014 21 ROXT [m/(kg)~l/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)~2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0379 QOUT 0.0003 24 QCOOL [ " " j 6.7965 QCOOL 40.1789 25 MRPRE [kg^l/3J 1.8124 MRPRE 10.7144 A 0 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 14991 5148 0 0 30 TIMEB [sec] 102.52 78.94 0.00 0.00 31 OPT. FEED SYSTEM pump pump n/a n/a 32 PT [MPa] 0.3447 0.3447 0,.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0.000934 0.000934 0.000934 0.000934 35 VF [m3/kg] 0.015592 0.015592 0.015592 0.015592 36 OXTC 0.0053 0.0041 0,.0000 0.0000 37 FTC 0.0159 0.0122 0,.0000 0.0000 38 FSC 0.0833 0.0833 0 .0833 0.0833 J 7 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3079 0.3030 0 .2866 0.2866 42 COEFTL 0.3079 0.3030 0 .2866 0.2866 43 COEFTLM 0.0000 0.0000 0 .0000 0.0000 44 NUMER 0.3473 0.4315 0,.7134 0.7134 45 DENOM 5.2249 4.0230 0,.0000 0.0000 46 VALUE 6.6466 10.7256 0,.0000 0.0000 47 VAL2 0.6921 0.6970 0,.7134 0.7134 48 V2ML 2405.33 1042.05 0.00 0.00 49 174 J AC AD AE AF AG AH 50 SOLVING CUBIC FOR MF 51 A1 -1.101E+03 -2.951E+02 0.OOOE+OO O.OOOE+OO 52 A2 3.929E+05 2.832E+04 O.OOOE+OO 0.000E+00 53 A3 -4.739E+07 -9.171E+05 0.000E+00 O.OOOE+OO 54 55 0 -3.669E+03 -2.342E+02 O.OOOE+OO O.OOOE+OO 56 R 1.017E+06 1.738E+04 O.OOOE+OO O.OOOE+OO 57 D 9.843E+11 2.894E+08 O.OOOE+OO O.OOOE+OO 58 S 126.1768 32.5211 0.0000 0.0000 59 T 29.0B13 7.2029 0.0000 0.0000 60 MF [kg] 522.18 138.08 0.00 0.00 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kg] 3475.31 1495.04 0.00 0.00 65 MG [kg] 391.53 116.63 0.40 0.40 66 MSS [kg] 208.60 61.99 0.00 0.00 67 MRF [kg] 704.15 300.69 0.00 0.00 68 MROX [kg] 370.62 152.68 0.00 0.00 69 70 ME [kg] 173.07 51.43 0.00 0.00 71 MSE [kg] 169.49 50.37 0.00 0.00 72 MFS [kg] 543.07 161.38 0.00 0.00 73 MN [kg] 383.19 113.87 0.00 0.00 74 MFT [kg] 103.69 23.72 0.00 0.00 75 MOXT [kg] 34.78 7.96 0.00 0.00 76 77 MF [kg] 522.18 138.08 0.00 0.00 78 MOX [kg] 2924.22 773.24 0.00 0.00 79 MP [kg] 3446.40 911.31 0.00 0.00 80 81 82 83 84 85 Ml [kg] 6557.50 2535.77 0.40 0.40 86 MO [kg] 9994.14 3432.23 0.00 0.00 87 175 AJ AK AL AM AN AO 1 Case 2 2 3 4 LEG 1 LEO to HOHMANN towards Mars 5 6 TOTAL DELTA V 3800 TOTAL # OF STAGES 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. Y6B FUEL REFRIG. VeB 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 27302.68 10157.06 0.00 0.00 12 DV [m/s] 1900 1900 0 0 13 ISFUSED [S] 445.95 445.95 445.95 445.95 14 MRATIO 0.6477 0.6477 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 1.1352 HF 0.0014 21 ROXT [m/(kg)~l/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)"2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0379 QOUT 0.0003 24 QCOOL [ " " j 6.7965 QCOOL 40.1789 25 MRPRE [kg"l/3] 1.8124 MRPRE 10.7144 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 100443 40283 0 0 30 TIMEB [see] 104.74 104.74 0.00 0.00 31 OPT. FEED SYSTEM pump pump n/a n/a 32 PT [MPa] 0.3447 0.3447 0.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0.000934 0.000934 0.000934 0.000934 35 VF [m3/kg] 0.015592 0.015592 0.015592 0.015592 36 OXTC 0.0055 0.0055 0.0000 0.0000 37 FTC 0.0162 0.0162 0.0000 0.0000 38 FSC 0.0833 0.0833 0.0833 0.0833 39 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3083 0.3083 0.2866 0. 2866 42 COEFTL 0.3083 0.3083 0.2866 0. 2866 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3394 0.3394 0.7134 0.7134 45 DENOM 5.3377 5.3377 0.0000 0.0000 46 VALUE 6.3580 6.3580 0.0000 0. 0000 47 VAL2 0.6917 0.6917 0.7134 0. 7134 48 V2ML 18884.17 7025.23 0.00 0.00 49 176 J AJ AK AL AM AN AO 50 SOLVING CUBIC FOR MF 51 A1 -8.928E+03 -3.332E+03 0.000E+00 O.OOOE+OO 52 A2 2.646E+07 3.663E+06 0.000E+00 O.OOOE+OO 53 A3 -2.620E+10 -1.349E+09 O.OOOE+OO O.OOOE+OO 54 55 Q -3.420E+04 -1.274E+04 O.OOOE+OO O.OOOE+OO 56 R 7.641E+07 1.064E+07 O.OOOE+OO O.OOOE+OO 57 D 5.798E+15 1.112E+14 O.OOOE+OO O.OOOE+OO 58 S 534.3221 276.7174 0.0000 0.0000 59 T 64.0068 46.0537 0.0000 0.0000 60 MF [kg] 3574.21 1433.46 0.00 0.00 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kg] 27302.68 10157.06 0.00 0.00 65 MG [kg] 2379.94 1002.30 0.40 0.40 66 MSS [kg] 1269.09 534.35 0.00 0.00 67 MRF [kg] 2509.63 1368.85 0.00 0.00 68 MROX [kg] 1336.08 726.61 0.00 0.00 69 70 ME [kg] 1052.95 443.34 0.00 0.00 71 MSE [kg] 1031.13 434.16 0.00 0.00 72 MFS [kg] 3303.93 1391.12 0.00 0.00 73 MN [kg] 2331.23 981.56 0.00 0.00 74 MFT [kg] 644 .42 271.33 0.00 0.00 75 MOXT [kg] 216.17 91.02 0.00 0.00 76 77 MF [kg] 3574.21 1433.46 0.00 0.00 78 MOX [kg] 20015.56 8027.38 0.00 0.00 79 MP [kg] 23589.77 9460.84 0.00 0.00 80 81 82 83 84 85 Ml [kg] 43377.25 17401.70 0.40 0.40 86 MO [kg] 66961.68 26855.44 0.00 0.00 87 177 J C 88 Case 2 89 90 91 MARS SURFACE DETAILS 92 93 MARS ROVER [kg] 150 94 OX. FACTORY [kg] 100.00 > MASS USED 100.00 95 FUEL FACTORY [kg] 1008.00 > MASS USED 0.00 96 97 OXIDIZER REFRIGERATION UNITS [kg] 98 STAGE I STAGE II STAGE III STAGE IV 99 LEG 6 6.1334 0.0000 0.0000 0.0000 100 LEG 5 11.4899 8.1817 0.0000 0.0000 101 LEG 4 37.1340 21.9007 0.0000 0.0000 102 103 TOTAL 84 .8397 MASS USED 0.0000 104 105 FUEL REFRIGERATION UNITS [kg] 106 STAGE I STAGE II STAGE III STAGE IV 107 LEG 6 16.2838 0.0000 0.0000 0.0000 108 LEG 5 30.6726 21.7818 0.0000 0.0000 109 LEG 4 99.7123 58.6890 0.0000 0.0000 110 111 TOTAL 227.1394 MASS USED 227.1394 112 113 OXIDIZER NECESSARY FOR RETURN [kg] 114 STAGE I STAGE II STAGE III STAGE IV 115 LEG 6 1.5186 0.0000 0.0000 0.0000 116 LEG 5 4.6372 2.5395 0.0000 0.0000 117 LEG 4 36.2066 14.4172 0.0000 0.0000 118 119 TOTAL 59.3192 MASS USED 59.3192 120 121 FUEL NECESSARY FOR RETURN [kg] 122 STAGE I STAGE II STAGE III STAGE IV 123 LEG 6 0.2712 0.0000 0.0000 0.0000 124 LEG 5 0.8281 0.4535 0.0000 0.0000 125 LEG 4 6.4655 2.5745 0.0000 0.0000 126 127 TOTAL 10.5927 > MASS USED 0.0000 128 J H I J 88 89 TOTAL MASS SUMMARY Case 2 90 91 PRE-MARS POST-MARS 92 MA [kg] 140 MA [kg] 0.6 93 MG [kg] 3949 MG [kg] 12.4 94 MSS [kg] 2104 MSS [kg] 4.0 95 MRF [kg] 5041 MRF [kg] 0.0 96 MROX [kg] 2659 MROX [kg] 0.0 97 98 ME [kg] 1746 ME [kg] 3.3 99 MSE [kg] 1709 MSE [kg] 3.3 100 MFS [kg] 5477 MFS [kg] 3.2 101 MN [kg] 3865 MN [kg] 7.4 102 MFT [kg] 1052 MFT [kg] 19.1 103 MOXT [kg] 353 MOXT [kg] 7.7 104 105 MF [kg] 5714 MF [kg] 10.6 106 MFL [kg] 427.9 MFL [kg] 2.1 107 MOX [kg] 31997 MOX [kg] 0.0 10B MOXL [kg] 239.6 MOXL [kg] 0.0 109 MP [kg] 37711 MP [kg] 69.9 110 111 ML [kg] 28095 ML [kg] 61.0 112 J O P R 88 Case 2 89 90 TOTAL MISSION 91 MA [kg] 140.35 MG [kg] 1748.88 92 MG [kg] 3961.B6 MSE [kg] 1712.65 93 MSS [kg] 2107.87 MFS [kg] 5480.26 94 MRF [kg] 5268.10 MN [kg] 3872.02 95 MROX [kg] 2659.20 MFT [kg] 1071.05 96 MOXT [kg] 360.57 97 MF [kg] 5724.37 98 MFL [kg] 430.03 99 MOX [kg] 31997.13 100 MOXL [kg] 239.63 MSAMPL [kg] 1.00 101 MP [kg] 37780.81 MSUPP [kg] 1.00 102 ML [kg] 28382.80 MROV [kg] 150.00 103 MOXFAC [kg] 100.00 104 MFFAC [kg] 0.00 105 106 107 MLEO [kg] 66356.30 108 180 C D E F l SHEET #11 Case 3 ALL EARTH TRANSPORTED CH4 AND ISR 2 MLEO [kg] 37885 3 4 LEG 6 HOHMANN to Highly eccentric Earth orbit 5 BBaBIIBBRIIIIBIll •••••••••••••••••••••••••••••••••laaaataaaMa 6 TOTAL DELTA V 1500 TOTAL # OF STAGES X 7 MODULAR ENGINES No AEROBRAKING Yes 8 OX. REFRIG. No FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 2.00 0.00 0.00 0.00 12 DV [m/s] 1500 0 0 0 13 ISPUSED [s] 356.05 356.05 356.05 356.05 14 MRATIO 0.6509 1.0000 1.0000 1.0000 15 RBS tAU] 1 1 " 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 0.2147 HF 0.5330 21 ROXT [m/(kg)Al/3] 0.0515 RFT 0.0719 22 QIN [W/(kgP2/3 ) 1 .7370 QIN 3.3832 23 QOUT [ " ") 0.0072 QOUT 0.0346 24 QCOOL [ " " ) 6.9194 QCOOL 13.3943 25 MRPRE [kg~l/3] 1.8452 MRPRE 3.5718 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 10 0 0 0 30 TIMEB [sec] 82.87 0.00 0.00 0.00 31 OPT. FEED SYSTEM press. n/a n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0.002293 0 .002293 0.002293 0.002293 35 VF [m3/kg] 0.006233 0,.006233 0.006233 0.006233 36 OXTC 0.0576 0.0000 0.0000 0.0000 37 FTC 0.0559 0.0000 0.0000 0.0000 38 FSC 0.0250 0.0250 0.0250 0.0250 J3 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3420 0.2285 0.2285 0.2285 42 COEFTL 0.4920 0.2285 0.2285 0.2285 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.1589 0.7715 0.7715 0.7715 45 DENOM 9.1877 0.0000 0.0000 0.0000 46 VALUE 1.7295 0.0000 0.0000 0.0000 47 VAL2 0.5080 0.7715 0.7715 0.7715 48 V2ML 1.0161 0.0000 0.0000 0.0000 49 181 A B C D E F SOLVING CUBIC FOR MF 51 A1 -1.762E+00 0.OOOE+OO O.OOOE+OO O.OOOE+OO 52 A2 1.035E+00 O.OOOE+OO 0.OOOE+OO O.OOOE+OO 53 A3 -2.028E-01 O.OOOE+OO O.OOOE+OO O.OOOE+OO 54 55 Q O.OOOE+OO 0.000E+00 .000E+00 O.OOOE+OO 56 R -3.212E-20 O.OOOE+OO .OOOE+OO 0.000E+00 57 D 1.032E-39 O.OOOE+OO •000E+00 O.OOOE+OO 58 S -0.0000 0.0000 0.0000 0.0000 59 T -0.0000 0.0000 0.0000 0.0000 60 MF [kg] 0.5875 0.0000 0.0000 0.0000 61 62 MASS SUMMARY (Hohmann to highly eccentric Earth orbit) 63 MA [kg] 0.66 0.00 0.00 0.00 64 ML [kg] 2.00 0.00 0.00 0.00 65 MG [kg] 0.66 0.40 0.40 0.40 66 MSS [kg] 0.14 0.00 0.00 0.00 67 MRF [kg] 0.00 0.00 0.00 0.00 68 MROX [kg] 0.00 0.00 0.00 0.00 69 70 ME [kg] 0.12 0.00 0.00 00 71 MSE [kg] 0.11 0.00 0.00 00 72 MFS [kg] 0.11 ,00 0.00 00 73 MN [kg] 0.26 ,00 0.00 00 74 MFT [kg] 0.25 00 0.00 00 75 MOXT [kg] 0.25 ,00 0.00 00 76 77 MF [kg] ,59 ,00 0.00 ,00 78 MOX [kg] 64 ,00 0.00 ,00 79 MP [kg] 23 ,00 0.00 ,00 80 81 NUMMOD 0 0 0 0 82 MMODTL [kg] 00 ,00 0.00 ,00 83 MODTHRTL [kg] 00 ,00 0.00 ,00 84 85 Ml [kg] 4.56 ,40 0.40 0.40 86 MO [kg] 6.39 ,00 0.00 0.00 87 182 K 1 Case 2 3 4 LEG 5 MPO to HOHMANN towards Earth 5 6 TOTAL DELTA V 2700 TOTAL * OF STAGES 7 MODULAR ENGINES NO AEROBRAKING No 8 OX. REFRIG. No FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 12.45 6.43 0.00 0.00 12 DV [m/s] 1350 1350 0 0 13 ISPUSED [s] 356.05 356.05 356.05 356.05 14 MRATIO 0.6794 0.6794 1.0000 1.0000 15 RBS [AU] 1 1 1 1 15 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 0.2147 HF 0.5330 21 ROXT [m/(kgri/3] 0.0515 RFT 0.0719 22 QIN [W/(kg)-2/3] 1.7370 QIN 3.3832 23 QOUT [ " " j 0.0072 QOUT 0.0346 24 QCOOL [ " " ] 6.9194 QCOOL 13.3943 25 MRPRE [kg~l/3J 1.8452 MRPRE 3.5718 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 36 19 0 0 30 TIMEB [sec] 76.09 76.09 0.00 0.00 31 OPT. FEED SYSTEM press. press. n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0.002293 0.002293 0.002293 0,.002293 35 VF [m3/kg] 0.006233 0.006233 0.006233 0,.006233 36 OXTC 0.0529 0.0529 0.0000 0.0000 37 FTC 0.0513 0.0513 0.0000 0.0000 38 FSC 0.0250 0.0250 0.0250 0.0250 J J 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3327 0.3327 0.2285 0.2285 42 COEFTL 0.3327 0.3327 0.2285 0.2285 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3467 0.3467 0.7715 0.7715 45 DENOM 8.4361 8.4361 0.0000 0.0000 46 VALUE 4.1102 4.1102 0.0000 0.0000 47 VAL2 0.6673 0.6673 0.7715 0.7715 48 V2ML 8.3078 4.2929 0.0000 0.0000 49 183 K H I J K M 50 SOLVING CUBIC FOR MF 51 A1 -6,.064E+00 -3. 133E+00 0.OOOE+OO O.OOOE+OO 52 A2 1,.226E+01 3.273E+00 O.OOOE+OO O.OOOE+OO 53 A3 -8,.258E+00 -1. 139E+00 0.OOOE+OO O.OOOE+OO 54 55 Q 0 .OOOE+OO 0. OOOE+OO OOOE+OO O.OOOE+OO 56 R 5,. 140E-19 -2.570E-19 OOOE+OO O.OOOE+OO 57 D 2 .642E-37 6.605E-38 OOOE+OO O.OOOE+OO 58 S 0.0000 0.0000 0.0000 0.0000 59 T -0.0000 -0.0000 0.0000 0.0000 60 MF [kg] 2.0212 1.0444 0.0000 0.0000 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 00 0.00 64 ML [kg] 12.45 6.43 00 0.00 65 MG [kg] 1.09 0.76 40 0.40 66 MSS [kg] 0.37 0.19 00 0.00 67 MRF [kg] 0.00 0.00 00 0.00 68 MROX [kg] 0.00 0.00 00 0.00 69 70 ME [kg] 0.31 0.16 0.00 00 71 MSE [kg] 0.30 0.15 0.00 00 72 MFS [kg] 0.29 0.15 0.00 00 73 MN [kg] 0.68 0.35 0.00 00 74 MFT [kg] 0.59 0.31 0.00 00 75 MOXT [kg] 0.61 0.31 0.00 00 76 77 MF [kg] 2.02 1.04 00 0.00 78 MOX [kg] 5.66 2.92 00 0.00 79 MP [kg] 7.68 3.97 00 0.00 80 81 NUMMOD 0 0 0 0 82 MMODTL [kg] 0.00 0.00 00 0.00 83 MODTHRTL [kg] 0.00 0.00 00 0.00 84 85 Ml [kg} 16.68 8.81 ,40 0.40 86 MO [kg] 23.96 12.38 00 0.00 184 K Q 1 Case 3 2 3 4 LEG 4 Mars Ascent Vehicle to MPO 5 6 TOTAL DELTA V 3500 TOTAL « OF STAGES 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. No FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 62.60 24 .09 0.00 0.00 12 DV [m/s] 1750 1750 0 0 13 ISPUSED [S] 356.05 356.05 356.05 356.05 14 MRATIO 0.6059 0.6059 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 0.2147 HF 0.5330 21 ROXT [m/(kg)~l/3] 0.0515 RFT 0.0719 22 QIN [W/(kg)~2/3] 1.7370 QIN 3.3832 23 QOUT [ " " ] 0.0072 QOUT 0.0346 24 QCOOL [ " " ] 6.9194 QCOOL 13.3943 25 MRPRE [kg^l/3] 1.8452 MRPRE 3.5718 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 242 93 0 0 30 TIMEB [sec] 93.54 93.54 0.00 0.00 31 OPT. FEED SYSTEM press. press. n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0.002293 0.002293 0 .002293 0,.002293 35 VF [m3/kg] 0.006233 0.006233 0,.006233 0,.006233 36 OXTC 0.0650 0.0650 0.0000 0.0000 37 FTC 0.0631 0.0631 0.0000 0.0000 38 FSC 0.0250 0.0250 0.0250 0.0250 7Q 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3566 0.3566 0.2285 0.2285 42 COEFTL 0.3566 0.3566 0.2285 0.2285 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.2493 0.2493 0.7715 0.7715 45 DENOM 10.3708 10.3708 0.0000 0.0000 46 VALUE 2.4041 2.4041 0.0000 0.0000 47 VAL2 0.6434 0.6434 0.7715 0.7715 48 V2ML 40.28 15.50 0.00 0.00 49 185 K 0 P Q R S 50 SOLVING CUBIC FOR MF 51 A1 -5.026E+01 -1.934E+01 0.OOOE+OO 0. OOOE+OO 52 A2 e.420E+02 1.247E+02 0.OOOE+OO 0. OOOE+OO 53 A3 -4.702E+03 -2.681E+02 0.OOOE+OO 0.OOOE+OO 54 55 Q -2.467E-17 1.234E-17 0•OOOE+OO 0.OOOE+OO 56 R 0.OOOE+OO -1.151E-16 0.OOOE+OO 0.OOOE+OO 57 D -1.502E-50 1.326E-32 0•OOOE+OO 0.OOOE+OO 58 S 0.0000 -0.0000 0.0000 0.0000 59 T -0.0000 -0.0000 0.0000 0.0000 60 MF [kg] 16.7534 6.4479 0.0000 0.0000 0£ 1 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kg] 62.60 24.09 0.00 0.00 65 MG [kg] 6.34 2.68 0.40 0.40 66 MSS [kg] 3.17 1.22 0.00 0.00 67 MRF [kg] 0.00 0.00 0.00 0.00 68 MROX [kg] 0.00 0.00 0.00 0.00 69 70 ME [kg] 2.63 1.01 0.00 00 71 MSE [kg] 2.57 0.99 0.00 00 72 MFS [kg] 2.47 0.95 0.00 00 73 MN [kg] 5.83 2.24 0.00 00 74 MFT [kg] 6.24 2.40 0.00 00 75 MOXT [kg] 6.43 2.48 0.00 00 76 77 MF [kg] 16 .75 6.45 0.00 0.00 78 MOX [kg] 46.91 18.05 0.00 0.00 79 MP [kg] 63.66 24.50 0.00 0.00 80 81 NUMMOD 0 0 0 0 82 MMODTL [kg] 0.00 0.00 0.00 ,00 83 MODTHRTL [kg] 0.00 0.00 0.00 ,00 84 85 Ml [kg] 98.28 38.07 0.40 .40 86 MO [kg] 161.54 62.17 0.00 ,00 87 186 X Y AA l Case 3 2 LEG 3 MPO to Mars Soft Landing 3 •aSBBSBSBSHKBKHMa •llllltllMallBBSIOIII maiBKB 4 TOTAL DELTA V 1000 TOTAL * OF STAGES 5 MODULAR ENGINES No AEROBRAKING Yes 6 OX. REFRIG. Yes FUEL REFRIG. NO 7 OX. SURF. REF. No FUEL SURF. REF. No e OX. ISRU Yes FUEL ISRU No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 337.46 0.00 0.00 0.00 12 DV [m/S] 1000 0 0 0 13 ISPUSED [s] 356.05 356.05 356.05 356.05 14 MRATIO 0.7510 1.0000 1.0000 1.0000 15 RBS [AU] 1 1 - 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX |W/m2] 1.1352 HF 2, 6879 21 ROXT [m/(kg)~l/3] 0.0515 RFT 0, 0719 22 QIN [W/(kg)'2/3] 1.7370 QIN 3. 3832 23 QOUT [ " " j 0.0379 QOUT 0. 1747 24 QCOOL [ " "J 6.7965 QCOOL 12. 8340 25 MRPRE [kg~l/3] 1.8124 MRPRE 3.4224 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 1156 0 0 0 30 TIMEB [sec] 59.10 0.00 0.00 0.00 31 OPT. FEED SYSTEM pump n/a n/a n/a 32 PT [MPa] 0.3447 0.3447 0.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0,.000934 0..000934 0..000934 0.000934 35 VF [m3/kg] 0..002524 0..002524 0,.002524 0 .002524 36 OXTC 0.0033 0.0000 0.0000 0.0000 37 FTC 0.0032 0.0000 0.0000 0.0000 38 FSC 0.0833 0.0833 0.0833 0.0833 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.2934 0.2868 0.2868 0.2868 42 COEFTL 0.4434 0.2868 0.2868 0.2868 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3077 0.7132 0.7132 0.7132 45 DENOM 6.5516 0.0000 0.0000 0.0000 46 VALUE 4.6962 0.0000 0.0000 0.0000 47 VAL2 0.5566 0.7132 0.7132 0.7132 48 V2ML 187.84 0.00 0.00 0.00 49 187 K V W X y z AA 50 SOLVING CUBIC FOR MF 51 A1 -1.204E+02 O.OOOE+OO 0.000E+00 0.000E+00 52 A2 4.800E+03 O.OOOE+OO O.OOOE+OO O.OOOE+OO 53 A3 -6.400E+04 O.OOOE+OO O.OOOE+OO O.OOOE+OO 54 55 Q -1.204E+01 O.OOOE+OO O.OOOE+OO O.OOOE+OO 56 R 3.632E+02 O.OOOE+OO O.OOOE+OO O.OOOE+OO 57 D 1.302E+05 O.OOOE+OO O.OOOE+OO O.OOOE+O0 5B S 8.9794 0.0000 0.0000 0.0000 59 T 1.3408 0.0000 0.0000 0.0000 60 HF [kg] 50.4696 0.0000 0.0000 0.0000 61 62 MASS SUMMARY 63 MA [kg] 64.93 0.00 0.00 0.00 64 ML [kg] 337.46 0.00 0.00 0.00 65 MG [kg] 26.37 0.40 0.40 0.40 66 MSS [kg] 13.85 0.00 0.00 0.00 67 MRF [kg] 0.00 0.00 0.00 0.00 68 MROX [kg] 49.17 0.00 0.00 0.00 69 70 ME [kg] 11.49 0.00 0.00 0.00 71 MSE [kg] 11.25 0.00 0.00 0.00 72 MFS [kg] 36.06 0.00 0.00 0.00 73 MN [kg] 25.51 0.00 0.00 0.00 74 MFT [kg] 1.40 0.00 0.00 0.00 75 MOXT [kg] 1.45 0.00 0.00 0.00 76 77 MF [kg] 50.47 0.00 0.00 0.00 78 MOX [kg] 141.31 0.00 0.00 0.00 79 MP [kg] 191.78 0.00 0.00 0.00 80 81 82 83 84 85 Ml [kg] 578.95 0.40 0.40 0.40 86 MO [kg] 770.34 0.00 0.00 0.00 87 188 AC AD AE AF AG AH Case 3 LEG 2 HOHMANN to Mars Parking Orbit (200 km) TOTAL DELTA V 2700 TOTAL # OF STAGES MODULAR ENGINES No AEROBRAKING No OX. REFRIG. Yes FUEL REFRIG. No 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 1682.64 773.67 0.00 0.00 12 DV [m/s] 1850 1350 0 0 13 ISPUSED [s] 356.05 356.05 356.05 356.05 14 MRATIO 0.5888 0.6794 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 7 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 1.1352 HF 2.6879 21 ROXT [m/(kgPl/3] 0.0515 RFT 0.0719 22 QIN [W/(kg)*2/3] 1.7370 QIN 3.3832 23 QOUT [ " " ] 0.0379 QOUT 0.1747 24 QCOOL [ " " ] 6.7965 QCOOL 12.8340 25 MRPRE [kg~l/3] 1.8124 MRPRE 3.4224 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 7298 2510 0 0 30 TIMEB [sec] 97.60 76.09 0.00 0.00 31 OPT. FEED SYSTEM pump pump n/a n/a 32 PT [MPa] 0.3447 0. 3447 0,,3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0.000934 0,.000934 0.000934 0..000934 35 7F [m3/kg] 0.002524 0 .002524 0.002524 0,.002524 36 OXTC 0.0055 0.0043 0..0000 0.0000 37 FTC 0.0053 0.0042 0,,0000 0.0000 38 FSC 0.0833 0.0833 0..0833 0.0833 J"*Q 7 4C INTERMEDIATE MASS CALCS 41 COEFSUM 0.2976 0.2953 0..2868 0.2868 42 COEFTL 0.2976 0.2953 0,.2868 0.2868 43 COEFTLM 0.0000 0.0000 0..0000 0.0000 44 NUMER 0.2912 0.3842 0..7132 0.7132 45 DENOM 10.8208 6 .4361 0..0000 0.0000 46 VALUE 2.6908 4 .5540 0,.0000 0.0000 47 VAL2 0 .7024 0.7047 0..7132 0.7132 48 V2ML 1181.81 545.24 0.00 0.00 49 AC AO AE AF AG SOLVING CUBIC FOR MF 51 A1 -1.320E+03 -3.597E+02 0.OOOE+OO 52 A2 5.787E+05 4.300E+04 0.OOOE+OO 53 A3 -8.472E+07 -1.716E+06 0.OOOE+OO 54 55 Q -7.021E+02 -3.947E+01 OOOE+OO 56 R 2.319E+05 3.552E+03 OOOE+OO 57 D 5.344E+10 1.256E+07 OOOE+OO 58 S 77.3664 19.2157 0.0000 59 T 9.0753 2.0543 0.0000 60 MF [kg] 526.45 141.16 0.00 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 64 ML [kg] 1682.64 773.67 0.00 65 MG [kg] 191.35 54.38 0.40 66 MSS [kg] 101.84 28.79 0.00 67 MRF [kg] 0.00 0.00 0.00 68 MROX (kg] 234.74 97.61 0.00 69 70 ME [kg] 84.50 23.89 00 71 MSE [kg] 82 .75 23.39 00 72 MFS [kg] 265.13 74.95 00 73 MN [kg] 187.54 53.02 00 74 MFT [kg] 16.97 74 00 75 MOXT [kg] 17.58 88 00 76 77 MF [kg] 526.45 141.16 00 78 MOX [kg] 1474.06 395.26 00 79 MP [kg] 2000.50 536.42 00 80 81 82 83 84 85 Ml [kg] 2865.04 1137.31 0.40 86 MO [kg] 4865.15 1673.33 0.00 87 190 K AJ AK AL AM AN AO 1 Case 3 2 3 4 LEG 1 LEO to HOHMANN towards Mars 5 6 TOTAL DELTA V 3800 TOTAL # OF STAGES 2 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. Yes FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 13968.83 4899.89 0.00 0.00 12 DV [m/s] 1900 1900 0 0 13 ISPUSED [S] 356.05 356.05 356.05 356.05 14 MRATIO 0.5804 0.5804 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/ra2] 1388 1388 1388 1388 17 13 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 1.1352 HF 2.6879 21 ROXT [m/(kg)~l/3] 0.0515 RFT 0.0719 22 QIN [W/(kg)"2/3] 1.7370 QIN 3.3832 23 QOUT [ " " j 0.0379 QOUT 0.1747 24 QCOOL [ " "J 6.7965 QCOOL 12.8340 25 MRPRE [kgAl/3] 1.8124 MRPRE 3.4224 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 57038 20802 0 0 30 TIMEB [sec] 99.59 99.59 0.00 0.00 31 OPT. FEED SYSTEM pump pump n/a n/a 32 PT [MPa] 0.3447 0.3447 0.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0.000934 0.000934 0.000934 0..000934 35 VF [m3/kg] 0.002524 0.002524 0.002524 0..002524 36 OXTC 0.0056 0.0056 0.0000 0.0000 37 FTC 0.0054 0.0054 0.0000 0.0000 38 FSC 0.0833 0.0833 0.0833 0.0833 39 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.2979 0.2979 0..2868 0..2868 42 COEFTL 0.2979 0.2979 0..2668 0..2868 43 COEFTLM 0.0000 0.0000 0..0000 0..0000 44 NUMER 0.2826 0.2826 0..7132 0..7132 45 DENOM 11.0410 11.0410 0..0000 0..0000 46 VALUE 2.5593 2.5593 0..0000 0..0000 47 VAL2 0.7021 0.7021 0..7132 0.,7132 48 V2ML 9808.00 3440.38 0.00 0.00 49 191 AJ AK AL AM AN AO SOLVING CUBIC FOR MF 51 A1 -1.150E+04 -4.036E+03 0.000E+00 O.OOOE+OO 52 A2 4.406E+07 5.421E+06 0.000E+00 O.OOOE+OO 53 A3 -5.628E+10 -2.429E+09 O.OOOE+OO O.OOOE+OO 54 55 Q -7.114E+03 -2.496E+03 0.000E+00 O.OOOE+OO 56 R 2.045E+07 2.519E+06 O.OOOE+OO O.OOOE+OO 57 D 4.160E+14 6.330E+12 O.OOOE+OO O.OOOE+OO 58 S 344.5444 171.3963 0.0000 0.0000 59 T 20.6480 14.5628 0.0000 0.0000 60 MF [kg] 4198.40 1531.15 0.00 0.00 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 00 0.00 64 ML [kg] 13968.83 4899.89 00 0.00 65 MG [kg] 1443.79 538.47 40 0.40 66 MSS [kg] 769.81 286.97 00 0.00 67 MRF [kg] 0.00 0.00 00 0.00 68 MROX [kg] 937.02 478.30 00 0.00 69 70 ME [kg] 638.70 238.10 00 00 71 MSE [kg] 625.47 233.16 00 00 72 MFS [kg] 2004.12 747.10 00 00 73 MN [kg] 1417.61 528.46 00 00 74 MFT [kg] 130.89 48.79 00 00 75 MOXT [kg] 135.62 50.56 00 00 76 77 MF [kg] 4198.40 1531.15 00 0.00 78 MOX [kg] 11755.51 4287.21 00 0.00 79 MP [kg] 15953.91 5818.36 00 0.00 80 81 82 83 84 85 Ml [kg] 22071.86 8049.82 0.40 0.40 86 MO [kg] 38025.37 13867.78 0.00 0.00 87 K A B C D E F 88 Case 3 89 90 91 MARS SURFACE DETAILS 92 •••••BBttMaaBBiaaiBBBBBBBaBBaiHiaaBaBaatiaiitaaaiaBaHiaiiaaHiBi 93 MARS ROVER [kg] 150 94 OX. FACTORY [kg] 100.00 > MASS USED 100.00 95 FUEL FACTORY [kg] 1008.00 > MASS USED 0.00 96 97 OXIDIZER REFRIGERATION UNITS [Kg] 98 STAGE I STAGE II STAGE III STAGE IV 99 LEG 6 100 LEG 5 101 LEG 4 102 103 TOTAL 0.0000 MASS USED 0.0000 104 105 FUEL REFRIGERATION UNITS [kg] 106 STAGE I STAGE II STAGE III STAGE IV 107 LEG 6 108 LEG 5 109 LEG 4 110 111 TOTAL 0.0000 > MASS USED 0.0000 112 113 OXIDIZER NECESSARY FOR RETURN [kg] 114 STAGE I STAGE II STAGE III STAGE IV 115 LEG 6 1.6450 0.0000 0.0000 0.0000 116 LEG 5 5.6595 2.9244 0.0000 0.0000 117 LEG 4 46.9096 18.0542 0.0000 0.0000 118 119 TOTAL 75.1927 MASS USED 75.1927 120 121 FUEL NECESSARY FOR RETURN [kg] 122 STAGE I STAGE II STAGE III STAGE IV 123 LEG 6 0.5875 0.0000 0.0000 0.0000 124 LEG 5 2.0212 1.0444 0.0000 0.0000 125 LEG 4 16.7534 6.4479 0.0000 0.0000 126 127 TOTAL 26.8545 > MASS USED 0.0000 128 STAGE I STAGE II STAGE III STAGE IV 129 LEG 6 0.0000 0.0000 0.0000 0.0000 130 LEG 5 0.0000 0.0000 0.0000 0.0000 131 LEG 4 0.0000 0.0000 0.0000 0.0000 132 133 TOTAL 0.0000 134 193 K J K L M 88 89 TOTAL MASS SUMMARY 3 90 91 PRE-MARS POST-MARS 92 MA [kg] 65 MA [kg] 0.7 93 MG [kg] 2257 MG [kg] 14.3 94 MSS [kg] 1201 MSS [kg] 5.1 95 MRF [kg] 0 MRF [kg] 0.0 96 MROX [kg] 1797 MROX [kg] 0.0 97 98 ME [kg] 997 ME [kg] 4.2 99 MSE [kg] 976 MSE [kg] 4.1 100 MFS [kg j 3127 MFS [kgj 4.0 101 MN [kg] 2212 MN [kg] 9.4 102 MFT [kg] 202 MFT [kg] 9.8 103 MOXT [kg] 209 MOXT [kg] 10.1 104 105 MF [kg] 6448 MF [kg] 26.9 106 MFL [kg] 22.5 MFL [kg] 0.3 107 MOX [kg] 18053 MOX [kg] 0.0 108 MOXL [kg] 126.0 MOXL [kg] 0.0 109 MP [kg] 24501 MP [kg] 102.0 110 111 ML [kg] 13043 ML [kg] 61.6 112 K 0 P Q R S T 88 Case 3 89 90 TOTAL MISSION 91 MA [kg] 65.59 ME [kg] 1000. 89 92 MG [kg] 2271.50 MSE [kg] 980. 16 93 MSS [kg] 1206.35 MFS [kg] 3131.,33 94 MRF [kg] 0.00 MN [kg] 2221.50 95 MROX [kg] 1796.85 MFT [kg] 211.,58 96 MOXT [kg] 219.,16 97 MF [kg] 6474.48 98 MFL [kg] 22.76 99 MOX [kg] 18053.35 100 MOXL [kg] 125.96 MSAMPL [kg] 1.00 101 MP [kg] 24603.03 MSUPP [kg] 1.00 102 ML [kg] 13104.91 MROV [kg] 150.00 103 MOXFAC [kg] 100.00 104 MFFAC [kg] 0.00 105 106 107 MLEO [kg] 37884.75 108 195 L A B C D E F 1 SHEET #12 case 4 LH2+LOX (out)/EARTH TRANSP. CH4+I 2 MLEO [kg] 51933 3 4 LEG 6 HOHMANN to Highly eccentric Earth orbit 5 BlIBBRRBBSBOaBBBBBBIIBatBIBIIIIBRiaRHIBIBBRKBIIBBIIIRRIISIIKBIKIBI 6 TOTAL DELTA V 1500 TOTAL * OF STAGES 1 7 MODULAR ENGINES No AEROBRAKING Yes 8 OX. REFRIG. No FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML (kg] 2.00 0.00 0.00 0.00 12 DV jra/s] 1500 0 0 0 13 ISPUSED [s] 356.05 356.05 356.05 356.05 14 MRATIO 0.6509 1.0000 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX (W/m2] 0.2147 HF 0.5330 21 ROXT [m/(kgri/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)~2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0072 QOUT 0.1029 24 QCOOL [ " " j 6.9194 QCOOL 39.7685 25 MRPRE [kg-1/3] 1.6452 MRPRE 10.6049 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 10 0 0 0 30 TIMEB [sec] 82.87 0.00 0.00 0.00 31 OPT. FEED SYSTEM press. n/a n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0.002293 0,.002293 0 .002293 0 .002293 35 VF [m3/kg] 0.006233 0..006233 0,.006233 0 .006233 36 OXTC 0.0576 0.0000 0.0000 0.0000 37 FTC 0.0559 0.0000 0.0000 0.0000 38 FSC 0.0250 0.0250 0.0250 0.0250 39 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3420 0.2285 0..2285 0.2285 42 COEFTL 0.4920 0.2285 0..2285 0.2285 43 COEFTLM 0.0000 0.0000 0..0000 0.0000 44 NUMER 0.1589 0.7715 0.,7715 0.7715 45 DENOM 9.1877 0.0000 0..0000 0.0000 46 VALUE 1.7295 0.0000 0..0000 0.0000 47 VAL2 0.5080 0.7715 0.,7715 0.7715 48 V2ML 1.0161 0.0000 0..0000 0.0000 49 196 L A B C D E F 50 SOLVING CUBIC FOR MF 51 A1 -1.762E+00 0.000E+00 O.OOOE+OO O.OOOE+OO 52 A2 1.035E+00 0.000E+00 0.000E+00 O.OOOE+OO 53 A3 -2.028E-01 O.OOOE+OO 0.000E-I-00 O.OOOE+OO 54 55 Q 0.000E+00 O.OOOE+OO O.OOOE+OO O.OOOE+OO 56 R -3.212E-20 O.OOOE+OO O.OOOE+OO O.OOOE+OO 57 D 1.032E-39 O.OOOE+OO O.OOOE+OO O.OOOE+OO 58 S -0.0000 0.0000 0.0000 0.0000 59 T -0.0000 0.0000 0.0000 0.0000 60 MF [kg] 0.5875 0.0000 0.0000 0.0000 61 62 MASS SUMMARY (Hohmann to highly eccentric Earth orbit) 63 MA [kg] 0.66 0.00 0.00 0.00 64 ML [kg] 2.00 0.00 0.00 0.00 65 MG [kg j 0.66 0.40 0.40 0.40 66 MSS [kg] 0.14 0.00 0.00 0.00 67 MRF [kg] 0.00 0.00 0.00 0.00 68 MROX [kg] 0.00 0.00 0.00 0.00 69 70 ME [kg] 0.12 0.00 0.00 0.00 71 MSE [kg] 0.11 0.00 0.00 0.00 72 MFS [kg] 0.11 0.00 0.00 0.00 73 MN [kg] 0.26 0.00 0.00 0.00 74 MFT [kg] 0.25 0.00 0.00 0.00 75 MOXT [kg] 0.25 0.00 0.00 0.00 76 77 MF [kg] 0.59 0.00 0.00 0.00 78 MOX [kg] 1.64 0.00 0.00 0.00 79 MP [kg] 2.23 0.00 0.00 0.00 80 81 NUMMOD 0 0 0 0 82 MMODTL [kg] 0.00 0.00 0.00 0.00 83 MODTHRTL [kg] 0.00 0.00 0.00 0.00 84 85 Ml [kg] 4.56 0.40 0.40 0.40 86 MO [kg] 6.39 0.00 0.00 0.00 87 197 L H J 1 OX (in) Case 4 2 3 4 LEG 5 MPO to HOHMANN towards Earth 5 6 TOTAL DELTA V 2700 TOTAL # OF STAGES 2 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. No FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 12.45 6.43 0.00 0.00 12 DV [m/s] 1350 1350 0 0 13 ISPUSED [s] 356.05 356.05 356.05 356.05 14 MRATIO 0.6794 0.6794 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 0.2147 HF 0.5330 21 ROXT [m/(kg)^l/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)A2/3] 1.7370 QIN 10.0450 23 QOUT [ " " j 0.0072 QOUT 0.1029 24 QCOOL [ " " ] 6.9194 QCOOL 39.7685 25 MRPRE [kg'1/3] 1.8452 MRPRE 10.6049 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 36 19 0 0 30 TIMEB [sec] 76.09 76.09 0.00 0.00 31 OPT. FEED SYSTEM press. press. n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0 .002293 0.002293 0.002293 0,.002293 35 VF [m3/kg] 0 .006233 0.006233 0.006233 0 .006233 36 OXTC 0.0529 0.0529 0.0000 0.0000 37 FTC 0.0513 0.0513 0.0000 0.0000 38 FSC 0.0250 0.0250 0.0250 0.0250 39 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3327 0.3327 0..2285 0.2285 42 COEFTL 0.3327 0.3327 0..2285 0.2285 43 COEFTLM 0.0000 0.0000 0..0000 0.0000 44 NUMER 0.3467 0.3467 0..7715 0.7715 45 DENOM 8.4361 8.4361 0..0000 0.0000 46 VALUE 4.1102 4.1102 0..0000 0.0000 47 VAL2 0.6673 0.6673 0,.7715 0.7715 48 V2ML 8.3078 4.2929 0,.0000 0.0000 49 198 H I M SOLVING CUBIC FOR MF 51 A1 • -6.064E+00 -3.133E+00 0.OOOE+OO 0.OOOE+OO 52 A2 1.226E+01 3.273E+00 0.000E+00 0.OOOE+OO 53 A3 -8.258E+00 -1.139E+00 0.000E+00 0.OOOE+OO 54 55 Q 0.OOOE+OO 000E+00 0.000E+00 0.OOOE+OO 56 R 5.140E-19 570E-19 0.OOOE+OO 0.OOOE+OO 57 D 2.642E-37 605E-38 0.000E+00 0.OOOE+OO 58 S 0.0000 0.0000 0.0000 0.0000 59 T -0.0000 -0.0000 0.0000 0.0000 60 MF [kg] 2.0212 1.0444 0.0000 0.0000 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kg] 12.45 6.43 0.00 0.00 65 MG [kg] 1.09 0.76 0.40 0.40 66 MSS [kg] 0.37 0.19 0.00 0.00 67 MRF [kg j 0.00 0.00 0.00 0.00 68 MROX [kg] 0.00 0.00 0.00 0.00 69 70 ME [kg] 0.31 0.16 0.00 00 71 MSE [kg] 0.30 0.15 0.00 00 72 MFS [kg] 0.29 0.15 0.00 00 73 MN [kg] 0.68 0.35 0.00 00 74 MFT [kg) 0.59 0.31 0.00 00 75 MOXT [kg] 0.61 0.31 0.00 00 76 77 MF [kg] 2.02 1.04 0.00 0.00 78 MOX [kg] 5.66 2.92 0.00 0.00 79 MP [kg] 7.68 3.97 0.00 0.00 80 81 NUMMOD 0 0 0 0 82 MMODTL [kg] 0.00 0.00 0.00 0.00 83 MODTHRTL [kg] 0.00 0.00 0.00 0.00 84 85 Ml [kg] 16.68 8.81 0.40 0.40 86 MO [kg] 23.96 12.38 0.00 0.00 87 199 L Q 1 Case 4 2 3 4 LEG 4 Mars Ascent Vehicle to MPO 5 6 TOTAL DELTA V 3500 TOTAL # OF STAGES 7 MODULAR ENGINES No AEROBRAKING No 8 OX. REFRIG. NO FUEL REFRIG. No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 62.60 24.09 0.00 0.00 12 DV [m/s] 1750 1750 0 0 13 ISPUSED [s] 356.05 356.05 356.05 356.05 14 MRATIO 0.6059 0.6059 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 IB REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2) 0.2147 HF 0.5330 21 ROXT [m/(kg)"l/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)"2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0072 QOUT 0.1029 24 QCOOL [ " " ] 6.9194 QCOOL 39.7685 25 MRPRE [kg"l/3] 1.8452 MRPRE 10.6049 26 27 ROCKET DESIGN DETAILS (stage dependence) 2B FEED SYSTEM USED press. press. press. press. 29 THRDES [kg] 242 93 0 0 30 TIMEB [sec] 93.54 93.54 0.00 0.00 31 OPT. FEED SYSTEM press. press. n/a n/a 32 PT [MPa] 5.1710 5.1710 5.1710 5.1710 33 PINDEX 7 7 7 7 34 VOX [m3/kg] 0 .002293 0 .002293 0.002293 0.002293 35 VF [m3/kg] 0 .006233 0 .006233 0.006233 0.006233 36 CXTC 0.0650 0.0650 0.0000 0.0000 37 FTC 0.0631 0.0631 0.0000 0.0000 38 FSC 0.0250 0.0250 0.0250 0.0250 39 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3566 0.3566 0..2285 0..2285 42 COEFTL 0.3566 0.3566 0..2285 0..2285 43 COEFTLM 0.0000 0.0000 0..0000 0..0000 44 NUMER 0.2493 0.2493 0..7715 0.,7715 45 DENOM 10.3708 10.3708 0..0000 0..0000 46 VALUE 2.4041 2.4041 0..0000 0..0000 47 VAL2 0.6434 0.6434 0..7715 0..7715 48 V2ML 40.28 15.50 0.00 0.00 49 200 L 0 P Q R SO SOLVING CUBIC FOR MF 51 A1 -5.026E+01 -1.. 934E+01 0.OOOE+OO 0.OOOE+OO 52 A2 8.420E+02 1,.247E+02 0.OOOE+OO 0.OOOE+OO 53 A3 -4.702E+03 -2,.681E+02 0.OOOE+OO 0.OOOE+OO 54 55 Q -2.467E-17 1..234E-17 0.OOOE+OO 0.OOOE+OO 56 R 0.OOOE+OO -1..151E-16 0.OOOE+OO 0.OOOE+OO 57 D -1.502E-50 1..326E-32 0.OOOE+OO 0.OOOE+OO 58 S 0.0000 -0.0000 0.0000 0.0000 59 T -0.0000 -0.0000 0.0000 0.0000 60 MF [kg] 16.7534 6.4479 0.0000 0.0000 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kg] 62.60 24.09 0.00 0.00 65 MG [kg] 6.34 2.68 0.40 0.40 66 MSS [kg] 3.17 1.22 0.00 0.00 67 MRF [kg] 0.00 0.00 0.00 0.00 6B MROX [kg] 0.00 0.00 0.00 0.00 69 70 ME [kg] 2.63 X.01 0.00 0.00 71 MSE [kg] 2.57 0.99 0.00 0.00 72 MFS [kg] 2.47 0.95 0.00 0.00 73 MN [kg] 5.82 2.24 0.00 0.00 74 MFT [kg] 6.24 2.40 0.00 0.00 75 MOXT [kg] 6.43 2.48 0.00 0.00 76 77 MF [kg] 16.75 6.45 0.00 0.00 78 MOX [kg] 46.91 18.05 0.00 0.00 79 MP [kg] 63.66 24.50 0.00 0.00 80 81 NUMMOD 0 0 0 0 82 MMODTL [kg] 0.00 0.00 0.00 0.00 83 MODTHRTL [kg] 0.00 0.00 0.00 0.00 84 85 Ml [kg] 98.27 38.07 0.40 0.40 86 MO [kg] 161.54 62.17 0.00 0.00 201 L x y AA 1 Case 4 2 LEG 3 MPO to Mars Soft Landing 3 4 TOTAL DELTA V 1000 TOTAL * OF STAGES 5 MODULAR ENGINES No AEROBRAXING Yes 6 OX. REFRIG. Yes FUEL REFRIG. Yes 7 OX. SURF. REF. No FUEL SURF. REF No 8 OX. ISRU Yes FUEL ISRU No 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 337.46 0.00 0.00 0.00 12 DV [m/s] 1000 0 0 0 13 ISPUSED [S] 445.95 445.95 445.95 445.95 14 MRATIO 0.7957 1.0000 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 1.1352 HF 0.0014 21 ROXT [m/(kg)"l/3] 0.0515 RFT 0.1239 22 QIN [W/(kgr2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.037 9 QOUT 0.0003 24 QCOOL [ " " j 6.7965 QCOOL 40.1789 25 MRPRE [kg~l/3] 1.8124 MRPRE 10.7144 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 1509 0 0 0 30 TIMEB (secj 60.75 0.00 0.00 0.00 31 OPT. FEED SYSTEM pump n/a n/a n/a 32 PT [MPa] 0.3447 0.3447 0.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0.000934 0,.000934 0.000934 0,.000934 35 VF [m3/kg] 0.015592 0..015592 0.015592 0..015592 36 OXTC 0.0032 0.0000 0.0000 0.0000 37 FTC 0.0094 0.0000 0.0000 0.0000 38 FSC 0.0833 0.0833 0.0833 0.0833 J 7 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.2992 0.2866 0.2866 0.2866 42 COEFTL 0.4492 0.2866 0.2866 0.2866 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3464 0.7134 0.7134 0.7134 45 DENOM 3.0961 0.0000 0.0000 0.0000 46 VALUE 11.1895 0.0000 0.0000 0.0000 47 VAL2 0.5508 0.7134 0.7134 0.7134 48 V2ML 185.86 0.00 0.00 0.00 49 L V W X Y Z AA 50 SOLVING CUBIC FOR MF 51 A1 -5.300E+01 O.OOOE+OO 0.OOOE+OO O.OOOE+OO 52 A2 8.277E+02 O.OOOE+OO 0.OOOE+OO O.OOOE+OO 53 A3 -4.583E+03 0.OOOE+OO O.OOOE+OO O.OOOE+OO 54 55 Q -3.617E+01 O.OOOE+OO O.OOOE+OO O.OOOE+OO 56 R 4.934E+02 O.OOOE+OO O.OOOE+OO O.OOOE+OO 57 D 1.961E+05 O.OOOE+OO O.OOOE+OO O.OOOE+OO 58 S 9.7828 0.0000 0.0000 0.0000 59 T 3.6973 0.0000 0.0000 0.0000 60 MF [kg] 31.1458 0.0000 0.0000 0.0000 61 62 MASS SUMMARY 63 MA [kg] 100.28 0.00 0.00 0.00 64 ML [kg] 337.46 0.00 0.00 0.00 65 MG [kg] 40.51 0.40 0.40 0.40 66 MSS [kg] 21.39 0.00 0.00 0.00 67 MRF [kg] 106.06 0.00 0.00 0.00 68 MROX [kg] 56.58 0.00 0.00 0.00 69 70 ME [kg] 17.75 0.00 0.00 0.00 71 MSE [kg] 17.38 0.00 0.00 0.00 72 MFS [kg] 55.69 0.00 0.00 0.00 73 MN [kg] 39.30 0.00 0.00 0.00 74 MFT [kg] 6.30 0.00 0.00 0.00 75 MOXT [kg] 2.11 0.00 0.00 0.00 76 77 MF [kg] 31.15 0.00 0.00 0.00 7 B MOX [kg] 174.42 0.00 0.00 0.00 79 MP [kg] 205.56 0.00 0.00 0.00 80 81 82 83 84 85 Ml [kg] 800.82 0.40 0.40 0.40 86 MO [kg] 1005.98 0.00 0.00 0.00 87 203 L AC AD AE AF AG AH 1 Case 4 2 3 4 LEG 2 HOHMANN to Mars Parking Orbit (200 km) 5 aMaiii«BlliaRngail»HtllR»aai IMIIIIIIIIBIBIIBRBIIMIIIItlMIMBB 6 TOTAL DELTA V 2700 TOTAL * OF STAGES 2 7 MODULAR ENGINES No AEROBRAKING No e OX. REFRIG. Yes FUEL REFRIG. Yes 9 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 2484.88 1015.70 0.00 0.00 12 DV [m/s] 1850 1350 0 0 13 ISPUSED [s] 445.95 445.95 445.95 445.95 14 MRATIO 0.6552 0.7345 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 17 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2) 1.1352 HF 0.0014 21 ROXT [m/(kg)*l/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)~2/3] 1.7370 QIN 10.0450 23 QOUT [ " » ] 0.0379 QOUT 0.0003 24 QCOOL I " " ] 6.7965 QCOOL 40.1789 25 MRPRE [kg~l/3] 1.8124 MRPRE 10.7144 26 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg] 11223 3681 0 0 30 TIMEB [see] 102.52 78.94 0.00 0.00 31 OPT. FEED SYSTEM pump pump n/a n/a 32 PT [MPa] 0.3447 0.3447 0.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0.000934 0.000934 0.000934 0,.000934 35 VF [m3/kg] 0.015592 0.015592 0.015592 0 .015592 36 OXTC 0.0053 0.0041 0.0000 0.0000 37 FTC 0.0159 0.0122 0.0000 0.0000 38 FSC 0.0833 0.0833 0.0833 0.0833 39 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3079 0.3030 0.2866 0.2866 42 COEFTL 0.3079 0.3030 0.2866 0.2866 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3473 0.4315 0.7134 0.7134 45 DENOM 5.2249 4.0230 0.0000 0.0000 46 VALUE 6.6466 10.7256 0.0000 0.0000 47 VAL2 0.6921 0.6970 0.7134 0.7134 48 V2ML 1719.83 707.95 0.00 0.00 49 204 L AC AD AE AF AG AH 50 SOLVING CUBIC FOR MF 51 A1 -7.914E+02 -2.016E+02 0.000E+00 0.000E+00 52 A2 2.009E+05 1.307E+04 O.OOOE+OO 0.000E+00 53 A3 -1.732E+07 -2.876E+05 0.000E+00 O.OOOE+OO 54 55 Q -2.631E+03 -1.596E+02 O.OOOE+OO O.OOOE+OO 56 R 5.254E+05 8.1X6E+03 O.OOOE+OO O.OOOE+OO 57 D 2.579E+11 6.180E+07 O.OOOE+OO O.OOOE+OO 58 S 101.0970 25.1865 0.0000 0.0000 59 T 26.0233 6.3368 0.0000 0.0000 60 MF [kg] 390.91 98.73 0.00 0.00 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kg] 2484.88 1015.70 0.00 0.00 65 MG [kg] 300.21 86.70 0.40 0.40 66 MSS [kg] 159.90 46.03 0.00 0.00 67 MRF [kg] 572.82 228.87 0.00 0.00 68 MROX [kg] 305.56 122.09 0.00 0.00 69 70 ME [kg] 132.67 38.19 0.00 0.00 71 MSE [kg] 129.92 37.40 0.00 0.00 72 MFS [kg] 416.28 119.83 0.00 0.00 73 MN [kg] 293.72 84.55 0.00 0.00 74 MFT [kg] 79.48 17.62 0.00 0.00 75 MOXT [kg] 26.66 5.91 0.00 0.00 76 77 MF [kg] 390.91 98.73 0.00 0.00 78 MOX [kg] 2189.09 552.87 0.00 0.00 79 MP [kg] 2580.00 651.60 0.00 0.00 80 81 82 83 84 85 Ml [kg] 4902.08 1802.88 0.40 0.40 86 MO [kg] 7481.68 2454.08 0.00 0.00 87 L AJ AK AL AM AN AO 1 Case 4 2 3 4 LEG 1 LEO to HOHMANN towards Mars 3e 6 TOTAL DELTA V 3800 TOTAL # OF STAGES 2 7 MODULAR ENGINES NO AEROBRAKING No B OX. REFRIG. Yes FUEL REFRIG. Yes Q 10 STAGE I STAGE II STAGE III STAGE IV 11 ML [kg] 21009.10 7603.65 0.00 0.00 12 DV [m/s] 1900 1900 0 0 13 ISPUSED [s] 445.95 445.95 445.95 445.95 14 MRATIO 0.6477 0.6477 1.0000 1.0000 15 RBS [AU] 1 1 1 1 16 SFB [W/m2] 1388 1388 1388 1388 1 7/ 18 REFRIGERATION MASS PRELIMINARY CALCULATIONS 19 Oxidizer Fuel 20 HOX [W/m2] 1.1352 HF 0.0014 21 ROXT [m/(kg)~l/3] 0.0515 RFT 0.1239 22 QIN [W/(kg)-2/3] 1.7370 QIN 10.0450 23 QOUT [ " " ] 0.0379 QOUT 0.0003 24 QCOOL [ " "J 6.7965 QCOOL 40.1789 25 MRPRE {kg"l/3] 1.8124 MRPRE 10.7144 27 ROCKET DESIGN DETAILS (stage dependence) 28 FEED SYSTEM USED pump pump pump pump 29 THRDES [kg) 78648 30997 0 0 30 TIMEB [see] 104.74 104.74 0.00 0.00 31 OPT. FEED SYSTEM pump pump n/a n/a 32 PT [MPa] 0.3447 0.3447 0.3447 0.3447 33 PINDEX 4 4 4 4 34 VOX [m3/kg] 0.000934 0.000934 0., 000934 0.000934 35 VF [m3/kg] 0.015592 0.015592 0. 015592 0.015592 36 OXTC 0.0055 0.0055 0.0000 0.0000 37 FTC 0.0162 0.0162 0.0000 0.0000 38 FSC 0.0833 0.0833 0.0833 0.0833 J3Q 7 40 INTERMEDIATE MASS CALCS 41 COEFSUM 0.3083 0.3083 0.2866 0.2866 42 COEFTL 0.3083 0.3083 0.2866 0.2866 43 COEFTLM 0.0000 0.0000 0.0000 0.0000 44 NUMER 0.3394 0.3394 0.7134 0.7134 45 DENOM 5.3377 5.3377 0.0000 0.0000 46 VALUE 6.3580 6.3580 0.0000 0.0000 47 VAL2 0.6917 0.6917 0.7134 0.7134 48 V2ML 14531.15 5259.14 0.00 0.00 206 L AJ AK AL AM AN AO 50 SOLVING CUBIC FOR MF 51 A1 -6.874E+03 -2.499E+03 O.OOOE+OO 0.000E+00 52 A2 1.567E+07 2.053E+06 O.OOOE+OO 0.000E+00 53 A3 -1.194E+10 -5.659E+08 O.OOOE+OO O.OOOE+OO 54 55 Q -2.632E+04 -9.548E+03 O.OOOE+OO O.OOOE+OO 56 R 4.529E+07 5.98SE+06 O.OOOE+OO O.OOOE+OO 57 D 2.033E+15 3.495E+13 O.OOOE+OO O.OOOE+OO 58 S 448.7776 228.2896 0.0000 0.0000 59 T 58.6579 41.8261 0.0000 0.0000 60 MF [kg] 2798.67 1103.03 0.00 0.00 61 62 MASS SUMMARY 63 MA [kg] 0.00 0.00 0.00 0.00 64 ML [kg] 21009.10 7603.65 0.00 0.00 65 MG [kg] 1885.78 784.08 0.40 0.40 66 MSS [kg] 1005.54 417.96 0.00 0.00 67 MRF [kg] 2127.82 1143.82 0.00 0.00 68 MROX [kg] 1135.04 610.15 0.00 0.00 69 70 ME [kg] 834.28 346.78 0.00 0.00 71 MSE [kg] 817.00 339.59 0.00 0.00 72 MFS [kg] 2617.81 1088.12 0.00 0.00 73 MN [kg] 1847.11 767.77 0.00 0.00 74 MFT [kg] 510.59 212.23 0.00 0.00 75 MOXT [kg] 171.28 71.19 0.00 0.00 76 77 MF [kg] 2798.67 1103.03 0.00 0.00 78 MOX [kg] 15672.53 6176.98 0.00 0.00 79 MP [kg] 18471.20 7280.01 0.00 0.00 BO 81 82 B3 84 85 Ml [kg] 33961.36 13385.35 0.40 0.40 B6 MO [kg] 52432.16 20664.96 0.00 0.00 B7 207 L A B C D E F 88 Case 4 89 90 91 MARS SURFACE DETAILS 92 BIBBBICIIBIBI11IIHIEBKIIBHHHSH •••••••••••••*•«••••• ••••••••••a 93 MARS ROVER [kg] 150 94 OX. FACTORY [kg] 100.00 > MASS USED 100.00 95 FUEL FACTORY [kg] 1008.00 > MASS USED 0.00 96 97 OXIDIZER REFRIGERATION UNITS [kg] 96 STAGE I STAGE II STAGE III STAGE IV 99 LEG 6 100 LEG 5 101 LEG 4 102 103 TOTAL 0.0000 > MASS USED 0.0000 104 105 FUEL REFRIGERATION UNITS [kg] 10S STAGE I STAGE II STAGE III STAGE IV 107 LEG 6 10B LEG 5 109 LEG 4 110 111 TOTAL 0.0000 > MASS USED 0.0000 112 113 OXIDIZER NECESSARY FOR RETURN [kg] 114 STAGE I STAGE II STAGE III STAGE IV 115 LEG 6 1.6450 0.0000 0.0000 0.0000 116 LEG 5 5.6595 2.9244 0.0000 0.0000 117 LEG 4 46.9096 18.0542 0.0000 0.0000 118 119 TOTAL 75.1927 > MASS USED 75.1927 120 121 FUEL NECESSARY FOR RETURN [kg] 122 STAGE I STAGE II STAGE III STAGE IV 123 LEG 6 0.5875 0.0000 0.0000 0.0000 124 LEG 5 2.0212 1.0444 0.0000 0.0000 125 LEG 4 16.7534 6.4479 0.0000 0.0000 126 127 TOTAL 26.8545 > MASS USED 0.0000 128 208 L I J K L M 88 89 , MASS SUMMARY Case 4 90 91 i-MARS POST-MARS 92 MA [kg] 100 MA [kg] 0.7 93 MG [kg] 3100 MG [kg] 14.3 94 MSS [kg] 1651 MSS [kg] 5.1 95 MRF [kg] 4179 MRF [kgj 0.0 96 MROX [kg] 2229 MROX [kg] 0.0 97 98 ME [kg] 1370 ME [kg] 4.2 99 MSE [kg] 1341 MSE [kg] 4.1 100 MFS [kg] 4298 MFS [kg] 4.0 101 MN [kg] 3032 MN [kg] 9.3 102 MFT [kg] 826 MFT [kg] 9.8 103 MOXT [kg] 277 MOXT [kg] 10.1 104 105 MF [kg] 4422 MF [kg] 26 .9 106 MFL [kg] 324.8 MFL [kg] 0.3 107 MOX [kg] 24766 MOX [kg] 0.0 108 MOXL [kg] 181.9 MOXL [kg] 0.0 109 MP [kg] 29188 MP [kg] 102.0 110 111 ML [kg] 22404 ML [kg] 61.6 112 L 0 P 0 R S T 88 Case 4 89 90 TOTAL MISSION 91 MA [kg] 100.94 ME [kg] 1373.88 92 MG [kg] 3114.42 MSE [kg] 1345.42 93 MSS [kg] 1655.90 MFS [kg] 4301.70 94 MRF [kg] 4179.39 MN [kg] 3041.78 95 MROX [kg] 2229.42 MFT [kg] 836.01 96 MOXT [kg] 287.24 97 MF [kg] 4449.34 98 MFL [kg] 325.03 99 MOX [kg] 24765.89 100 MOXL [kg] 181.87 MSAMPL [kg] 1.00 101 MP [kg] 29290.42 MSUPP [kg] 1.00 102 ML [kg] 22466.10 MROV [kg] 150.00 103 MOXFAC [kg] 100.00 104 MFFAC [kg] 0.00 105 106 107 MLEO [kg] 51933.33 108 210 A B C SHEET #13 - SPREADSHEET MAP OUTLINE OF SPECIFIC SHEETS 4 •EBeBBaBBBBBBSB 0 5 SHEET #1 Calculation Summary 6 SHEET *2 Mission Details 7 SHEET #3 Rocket Performance Data 8 SHEET #4 Rocket Propellant Data 9 SHEET #5 ISRU Component Data 10 SHEET #6 Support Components 11 SHEET #7 Spacecraft Details 12 SHEET #8 Modular Engine Details 13 SHEET #9 Case 1 ALL EARTH TRANSPORTED LH2 AND LOX 14 SHEET #10 Case 2 ALL EARTH TRANSPORTED LH2 AND ISRU LOX 15 SHEET #11 Case 3 ALL EARTH TRANSPORTED CH4 AND ISRU LOX 16 SHEET #12 CaBe 4 LH2+LOX (out)/EARTH TRANSP. CH4+ISRU LOX 17 SHEET #13 Spreadsheet Map is 19 20 (Sheets #9-12, Detailed map below) 21 22 Init. Page Tab 1 Tab 2 Tab 4 Tab 5 Tab 6 23 24 Leg 6 Leg 5 Leg 4 Leg 3 Leg 2 Leg 1 25 26 27 Pg-Dn 4 Pg-Dn 4 Pg-Dn 4 28 29 Surface Pre-Mars Total 30 Details Post-Mars Summary 31 Summary 32 211 M A B C 33 VARIABLE LIST 34 35 36 VARIABLE UNITS IDENTIFICATION 3 / 38 AC Aeroshell coefficient 39 AU AU Relative distance from sun 40 DF kg/m3 Fuel Density 41 DOX kg/m3 Oxidizer Density 42 EMISSF Fuel Emissivity 43 EMISSOX Oxidizer Emissivity 44 EC Engine Coefficient 45 FAF Fractional Area of Fuel exposed to sun 46 FAOX Fractional Area of Oxidizer exposed to s 47 FFL Fraction of Fuel Leaked 48 FOXL Fraction of Oxidizer Leaked 49 FSC Feed System Coefficient 50 FSF Fractional Scaling for Fuel emissivity 51 FSOX Fractional Scaling for Oxidizer emlssivi 52 FTC Fuel Tank Coefficient 53 FTR Fuel Tank Reflectivity 54 GC Guidance Coefficient 55 GEEZ Gravitational units (g's) 56 GZERO m/s2 Gravitational acceleration on Earth's Su 57 ISP sec Specific Impulse 58 MFFAC kg Fuel Factory Mass / 10 kg H2 / day 59 MOXFAC kg Oxygen Factory Mass / 10 kg 02 / day 60 MROV kg Rover Mass 61 MRPF kg/W Fuel Refrigeration Unit Specific Weight 62 MRPOX kg/W Oxidizer Refrigeration Unit Specific Wei 63 MSAMPL kg Sample Mass 64 MSUPP kg Support Mass 65 NC Nozzle Coefficient 66 NFT Number of Fuel Tanks 67 NCXT Number of Oxidizer Tanks 68 NRF Fuel Refrigerator Efficiency 69 NROX Oxidizer Refrigerator Efficiency 70 OXTC Oxidizer Tank Coefficient 71 OXTR Oxidizer Tank Reflectivity 72 PF % Fuel Percentage by maBS 73 PSTOR Pa Storage Pressure 74 SBC W/m2K4 Stefan-Boltzmann Constant 75 SCE Structural Coefficient - Engine 76 SCS Structural Coefficient - Surface Structu 77 SFE W/m2 Solar Flux in low Earth orbit 78 TF K Fuel Temperature 79 TOX K Oxidizer Temperature 80 THRUSTF Thrust Factor 81 212 M A B C D E F G 82 NSF Nozzle Safety Factor 63 DNOZ kg/m3 Density of Nozzle material 84 WNOZ m Average Width of Nozzle Material 85 HALFANGL deg Nozzle Expansion Half Angle 86 FTSFPRESS Fuel Tank design Safety Factor (Pressure 87 FTSFPUMP Fuel Tank design Safety Factor (Pump Fee 88 DFT kg/m3 Density of Fuel Tank material 89 SIGFT N/m2 Design stress of Fuel Tank material 90 OXTSFPRESS Oxygen Tank design Safety Factor (Press. 91 OXTSFPUMP Oxygen Tank design Safety Factor (Pump F 92 DOXT kg/m3 Density of Oxygen Tank material 93 SIGOXT N/m2 Design stress of Oxygen Tank material 94 95 STC Total Structural Coefficient 96 ML kg Payload Mass 97 DV m/s Delta V 98 1SPUSED sec Specific Impulse Used 99 MRATIO Propulsion equation Mass Ratio (exp[-DV/ 100 RBS AU Reference Distance: Body to Sun 101 SFB W/m2 Solar Flux at the Body 102 103 HOX W/m2 Oxidizer Heat Flux per unit area 104 ROXT m/kg^l/3 Radius of Oxygen Tank/(Oxidizer Mass)"(l 105 106 HF W/m2 Fuel Heat Flux per unit area 107 RFT m/kg^l/3 Radius of Fuel Tank/(Fuel Mass)~(l/3) 108 QIN m/kg~2/3 Heat input to tanks/(mass)~(2/3) 109 QOUT m/kg'k2/3 Heat output from tanks/(mass)'*(2/3) 110 QCOOL m/kg~2/3 Cooling rate for constant tank temperatu 111 MRPRE kg~(1/3) Preliminary calc. for refrigeration mass 112 113 THRDES kg Design Thrust 114 TIMEB sec Burn Time 115 OPT. FS Optimum type of feed system to use 116 PT MPa Tank Pressure 117 VOX m~3/kg Oxidizer Specific Volume 118 VF m~3/kg Fuel Specific Volume 119 OXTC Oxidizer Tank Coefficient 120 FTC Fuel Tank Coefficient 121 FSC Feed System Coefficient 122 123 COEFSUM Sum of coeff., excluding the aeroshell c 124 COEFTL Total sum of Coefficients 125 COEFTLM Modular engine Coefficient Total 126 NUMER Mass Ratio minus Total Sum of Coeff. 127 DENOM Fuel percentage * (1 - Mass Ratio) 128 VALUE 100 * Numerator / Denominator 129 VAL2 1 minus Total Sum of coefficients 130 V2ML kg Val2 * Payload Mass 131 213 M A B C D E F G 132 A1 Cubic equation coefficient 133 A2 Cubic equation coefficient 134 A3 Cubic equation coefficient 135 136 Q Eq. MF*3+A1*MFA2+A2*MF+A3«0 137 R Q,R,S,T are all intermediate calculation 138 S [See Schaum's Mathematical Handbook, 139 T Solutions of algebraic equations] 140 D Determinant 141 MF Real root for cubic solution 142 143 MA kg Aeroshell Mass 144 ML kg Payload Mass 145 MG kg Guidance Mass 146 MSS kg Structural Mass - Secondary 147 MRF kg Fuel Refrigeration Mass 148 MROX kg Oxidizer Refrigeration Mass 149 150 ME kg Engine Mass 151 MSE kg Structural Mass - Engine 152 MFS kg Feed System Mass 153 MN kg Nozzle Mass 154 MFT kg Fuel Tank Mass 155 MOXT kg Oxidizer Tank Mass 157 MF kg Fuel Mass 158 MOX kg Oxidizer Mass 159 MP kg Propellent mass (Mox+Mf) 160 161 NUMMOD Number of Modular Engines 162 MMODTL kg Total Mass of Modular Engines with prope 163 MODTHRTL kg Total Thrust of Modular Engines 164 165 Ml kg Final Mass 166 KO kg Initial Mass 167 APPENDIX B SPREADSHEET EQUATIONS 215 A2 [Will 'MARS SAMPLE RETURN MISSION ANALYSIS E2 [will 'Bruce Preiss 6/17/91 A3 (win B3 [Will C3 [Will D3 [Will £3 twin F3 [W12J A4 [Wll] '(Total sheets • 13. review sheet 13 for aap and additional info.) A7 [Will 'SHEET in - PROGRAH SUMMARY A9 [Wll] 'MASS SUMMARY FOR EACH MISSION Cll [Wll +5C:SAS2 D11 [Wll + SC:SAS22 Ell [Wll +SC:SAS42 Fll [W12 +SC:$AS62 C12 [Wll 1 IIIIIIIB D12 [Wll 1 BSBSSBCB E12 [Wll ' sssesssa F12 [W12 A1 3 [Wll 'FUEL (outbound) CI 3 [Wll +SC:SDS5 D1 3 [Wll +SC:SDS25 E13 [Wll + SC:$D$45 F13 [ W12 +$C:SDS65 A14 [Wll 'OXIDIZER (outbnd) Cl« [Wll +SC:SDS6 Dl« [Wll +SC:SDS 26 E14 [Wll +SC:$DS<<6 F14 [W12 +SC:$D$66 A15 [Wll •ISPP FUEL C15 [Wll #IF(SC:SGS1« "Yes" ,C17, D15 [Wll @IF(SC:SGS34 "Yes",D17, E15 (Wll gIF(SC:SGS54 "Yes",E17, F15 [W1 2 £IF(SC :SGS7A "Yes",F17, A16 [Wll 'ISPP OXID. C16 [Wll ^IF(SC:SGS15 "Yes",C18, D16 [Wll piF(SC:SGS35 "Yes",D18, E16 [Wll £IF($C:SGSS5 "Yes".E1S. F16 [ W1 2 {HF($C:$GS75="Yes",FlB, A17 [Wll 'FUEL (return) C17 [Wll + SC:S D$ 14 D17 [Wll + SC:S DS 34 E17 [Wll + SC:$D$5 A D21 [Villi 'Case 2 A E21 (Wll) "Case 3 A F21 [f 12] "Case 4 A C22 [Wll] \" A D22 [Wll] \- A E22 [Wll] \" A F22 [ W12] \- A B23 [Wll] 'MA (kg) A C23 (F0) [Will +SI:SQS91 A D23 (FO) [Wll] + SJ:SQS91 A E23 (FO) [Wll] + SK:SQS91 A F23 (FO) IW12] + $L:SQS91 A B24 [Wll] •MF [kg] A C24 (FO) [Wll] + SI SQS97 A D2« (FO) [Wll] + SJ:SQS97 A E24 (FO) [Wll] + SK: SQS97 A F24 (FO) [W12 ] + SL:SQS97 A B25 [Wll] 'MFL (kg) A C25 (FO) [Wll) + SI: SQS98 A D2S (FO) [Wll] + SJ:SQS98 A E25 (FO) [Will + SK: SQS98 A F25 (FO) [W12] + SL: SQS98 A B26 [Ull] •MFT [kg) A C26 (FO) (Wll) + SI STS95 A D26 (FO) [Wll] + SJ STS95 A E26 (FO) (Wll) + SK:STS95 A F26 (FO) [W12 ] + SL STS95 A B27 (Wll) 'MG (kg) A C27 (FO) [Wll] •SI SQS92 A D27 (FO) [Wll] + SJ: SQS92 A E27 (FO) [Wll] + SK:SQS92 A F27 (FO) [W12) •SL:SQS92 A B28 (Wll) ' HFS [kg] A C28 (FO) (Wll) •SI STS93 A D28 (FO) [Wll] + SJ:ST$93 A E28 (FO) [Wll] + SK:STS93 A F28 (FO) [W12] + SL:STS93 A B29 [Wll] •ME [kg) A C29 (FO) [Wll] •SI STS91 A D29 (FO) [Wll] + SJ: STS91 A E29 (FO) [Will + SK: STS91 A F29 (FO) [W12] + SL:STS91 A B30 [Wll] •MN (kg) A C 30 (FO) [Wll] •Si: STS94 A D30 (FO) [Wll] + SJ STS94 A E3C (FO) [Wll] +SK:STS94 A F30 (FO) [W12] + SL:STS94 A B31 (Wll] 'MNYFAC [kg] A C31 (FO) [Wll] •SI: STS104 A D31 (FO) [Wll] + SJ:STS104 A E31 (FO) [Wll] •SK: ST$104 A F31 (FO) [W12] + SL:STS104 A B32 (Wll] 'MOX [kg] A C32 (FO) [Wll] •Si: SQS99 A D32 (FO) (Wll) + SJ:SQS99 A E32 (FO [Nil] +$K:SQS99 A F32 (FO [W12] + SL:SQS99 A B33 [W1 ] 'MOXFAC [kg] A C33 (FO [Ml] +SI:STS103 A D33 (FO (Wll] +SJ:STS103 A E33 (FO [Nil] +$K:ST$103 A F33 (FO [W12] +$L:STS103 A B34 [W1 ] 'MOXL [kg) A C34 (FO [Wll] +$I:$Q$100 A D34 (FO [Will +$J:SQS100 A E34 (FO [»111 +SK:SQ$100 A F34 (FO [W12] +SL:SQS100 A B33 [W1 ] 'MOXT [kg] A C35 (FO [Wll] +SI:STS96 A D35 (FO [*11] +$J:STS96 A ESS (FO [Wll] +SK:STS96 A F35 (FO [W12] +SL:STS96 A B36 [W1 ) ' MRF [kg] A C36 (FO [Wll] +51:SQS94 A D36 (FO [Wll] + SJ: SQS94 A E36 (FO [Wll] + SK: SQS94 A F36 (FO [W12] +SL:SQS9^ A B37 (HI ] 'MROV [kg] A C37 (FO [Wll] +SI:STS102 A D37 (FO [Wll] +SJ:STS102 A E37 (FO [Wll] +SK:STS102 A F37 (FO [W12] + SL:STS102 A B36 (VI ] 'MROX [kg] A C30 (FO [Wll] +SI:SQS95 A D38 (FO [Wll] +SJ:SQS95 A E38 (FO [Wll] + SK:SQS95 A F38 (FO [ W12] +SL:SQS95 A B39 |W1 ] 'MS [kg] A C39 (FO [Wll] +SI:SQS93+SI STS92 A D39 (FO [Wll] •SJ:SQS93+SJ STS92 A E39 (FO [Wll] + SK:SQS93+SK STS92 A F39 (FO [W12] +SL:SQS93+SL STS92 A B40 [W1 ] 'MSAMPl [kg] A C<<0 (FO [Wll] + SIi STS100 A D40 (FO [Wll] + S J:STS100 A E40 (FO [Wll] + SK:STS100 A F<<0 (FO [W12] •SL:STS100 A BA1 (W1 ] 'MSUPP [kg] A cu\ (FO [Wll] +SI:STS101 A DM (FO [Wll] •SJ:STS101 A E41 (FO [Wll] +SK:STS101 A F41 (FO [W12] •SL:STS101 A C A F43 (F0) [W12] +SL:ST$107 A BA5 (Wll J *RF« INV RISK,REL.. A C45 (F2) [Wll] 0.4 A DA 5 (F2) [Wll] 0.3 A EA5 (F2) [Wll] 0.59 A FA5 (F2) [W12] 0.37 A BA6 (FO) [Wll] \- A CA6 (FO) [Wll] \- A DA6 (FO) [Wll] \- A EA6 (FO) [Wll] A FA6 (FO) [W12] \- A BA7 (Wll ] 'FoM A CA7 (FA) [Wll] (CAO+CA1)"10000"CA5/CA3 A DA7 (FA) [Wll] (DA0+DA1)"10000•DA5/DA3 A EA7 (FA) [Wll] (EAO+EA1)•10000•EA5/EA3 A FA7 (FA) [W12] (FAO+FA1)"10000•F45/F43 A BA8 (FO) [Wll] \- A CAB (FO) [Wll] \- A DA8 (FO) [Wll] \- A EA8 (FO) [Wll] \- A FA0 (FO) [ W12] \- A CA9 [Wll ] "Case 1 A DA9 (FA) [Wll) "Case 2 A EA9 (FA) (Wll) "Case 3 A FA9 (FA) (W12) "Case 4 B A1: 1*5) •SHEET #2 - MISSION DETAILS B A3: [W5] •PHYSICAL CONSTANTS B B5: [W9] •RADIAL DISTANCE FROM BODY TO SUN, EARTH B B6: [H9) "EARTH » B C6: [H9] 1 B D6: IV1B) 'AU B E6: [H6J "MARS » B F6: [«11] 1.324 B G6: [W9J •AU B B6: [H9] •MEAN RADIAL DISTANCE FROM BODY TO SUN [ B B9: [H9] "EARTH » B C9: (S2) [W9] 149599650000 B D9: (S2) [H18] '• B E9: (S2) (H6) "MARS - B F9: (S2) [H11] +C9-F6 B G9: [W9] 'm B Bll [W9] •GRAVITATIONAL ACCELERATION ON THE BODY B B12 [W9] "GZERO « B C12 (F4) [W 9] 9.81 B D12 (F4) [H18] 'm/s2 B E12 (F4) [W6J "GOMARS « B F12 (Fi.) [«11] 3.73 B G12 1W9] 'm/s2 B A13 [W5) 1 B A13 [W5] 'PROPULSIVE VELOCITY REQUIREMENTS B G15 [H9] *n/s B G1 6 |W9] \" B B17 (W9) 'LEG 1 B C17 (W9) 'LEO to HOHMANN towards Mars B G17 |H9) 3800 B CIS [ W9) '(No MID-COURSE CORRECTION used) B G18 IW91 0 B D19 (wie I 'Number of stages * B E19 [W6] 2 B D20 [Hie] 'Aerobraking * B E20 (H6 ] " S'o B D21 (Hie 1 'Modular engines • B E21 (H6] "No B B23 1H9) 'LEG 2 B C23 [H9] 'HOHMANN to Mars Parking Orbit (200 k«) B G23 [H9] 2700 B C24 (H9) '(includes a 0.3 [k«/s] MID-COURSE CORR B G24 [H9] 500 B D2S [H18 1 'Nunber of stages • B E25 [H6] 2 B D26 [Hie 1 'Aerobraking • B E26 [H6] "No B D27 [Hie 'Modular engines = B E27 [H6] "No B B29 [H9 J 'LEG 3 B C29 [H9] 1MPO to Mars Soft Landing B G29 [H9] 1000 B C30 [H9 ] '(No MID-COURSE CORRECTION used) B G30 [H9] 0 B D31 [HI 8 I 'Nunber of stages • 220 B E31 (W6 1 B 032 (VI ] 'Aerobraking • B E32 (W6 "Yes B D33 [W1 ] 'Modular engines * B E33 [H6 "No B B35 (W9 •LEG A B C35 (FA [H9] 'Mars Ascent Vehicle to MPO B G35 [W9 3500 B C36 [W9 '(No MID-COURSE CORRECTION used) B G36 [H9 0 B D37 (FA [WIS] 'Nuaber of stages • B E37 (FO [H6] 2 B D38 (HI ) 'Aerobraking * B E3B [W6 "No B D39 [Ml ] 'Modular engines » B E39 [W6 "No B B c A1: [W7] 'SHEET 03 - ROCKET PERFORMANCE DATA c A2: [W7 J 'Case 1 c C2: [wa] 'ALL EARTH TRANSPORTED LH2 AND LOX c E4: [Will "(Refrig.) c A5: [W7) ' Fuel (Outbound) c D5: [W8J "H2 c ES : (Wll) 'Yes c A6: [W7] ' Oxidizer(Outbound) c D6: [W8J "02 c E6: [Wll] "Yes c C7: (W8) 'Oxidizer/Fuel Ratio c F7: LWIS} 5.6 c G7: [W13] 'kg OX/kg fuel c C8: tW81 'Chamber Pressure c F8: (WIS)- 500 c Gfl : [W 1 3 j 'psia c C9: 1W8) 'Nozzle Expansion Ratio c F9: [W15] 50 c CIO (F2) {weJ 'Specific Impulse c FIO (WIS) 445.95 c GIO [W13] 'sec c CI 1 [W8] 'Characteristic Velocity c Fll (WIS) 7555 c Gil [ W13] 'ft/sec c El 3 IWll] "[Refrig.1 c F13 [W15] "[Surf. Refrig.) c G13 1W13] "I1SRU) c A14 [W 7| ' Fuel (Return) c D14 [W8] "H2 c E14 [Wll] "No c F14 (W15) -Yes c G14 [W13) 'No c A15 IW71 • Oxidizer (Return) c D15 [We] "02 c El 5 [Wll] -No c F1 5 [W15 J "Yes c G15 (W13) -No c C16 [W8] 'Oxidizer/Fuel Ratio c F16 (WIS) 5.6 c G16 (W13) 'kg OX/kg fuel c CI 7 |W8] 'Chaaber Pressure c F17 [WIS) 500 c G17 [W13) 'psia c C18 (W8] 'Nozzle Expansion Ratio c F18 (WIS) 50 c C19 (F2) [W6] 'Specific Impulse c F19 [WIS] 445.95 c G19 [W13] 'sec c C20 (W8] 'Characteristic Velocity c F20 [W15] 7555 c G20 [W13] 'ft/sec c A22 [W7] 'Case 2 c C22 (W8) 'ALL EARTH TRANSPORTED LH2 AND ISRU c E24 [Wll] "[Refrig.] c A25 [W7] ' Fuel (Outbound) 223 D2S [WBJ "H2 E25 [Wll "Yes A26 [W7] ' Oxidizer(Outbound) D26 [W8] "02 E26 [Wll "Yes C 27 1*8] 'Oxidizer/Fuel Ratio F27 [W15 5.6 G27 (W13 •kg OX/kg fuel C28 [WSJ 'Chamber Pressure F2B IW15 500 G28 (W1 3 'psia C29 [waj 'Nozzle Expansion Ratio F29 [W15 50 C30 [W8) 'Specific Impulse F30 [W15 <445.95 G30 [WIS 'sec C31 [W8] 'Characteristic Velocity F31 IW15 7555 G31 [ W1 3 'ft/sec E33 [Wll "[Refrig.] F33 [W15 "[Surf. Refrig.] G33 [WIS -[1SRU] AD't [ W7 ] Fuel (Return) D3<. [we] "H2 E34 [Wll "No F3A [W15 "Yes G3 c CU7 [W8] 'Oxidizer/Fuel Katio c F47 [M15] 2.8 c om (W1 3) 'kg OX/kg fuel c CUB [W8] 'Chamber Pressure c TUB [W15] 500 c OUB [W13] 'psia c C49 [W8] 'Nozzle Expansion Katio c F49 [W15] 50 c C50 [W8] 'Specific Iapulse c F50 [W15] 356.05 c G50 [HI 3] 1 sec c C51 [W8] 'Characteristic Velocity c F51 [WIS] 60*7 c G51 [W13] 'ft/sec c E53 [Wll] "[Re frig.] c F53 [W15] "[Surf. Refrig.] c G53 {W1 3) •[ISRU] c A54 IW7] Fuel (Return) c D5<. (W8] "CH* c E5<< [Will "No c F54 [WIS] "No c G5<. I w 1 3] -No c A55 [K7] Oxidizer (Return) c D55 [we ] "02 c E55 [Wll] •No c F55 H" 1 5 ] "No c G55 [ W1 3] "Yes c C56 [W8] 'Oxidizer/Fuel Ratio c F56 [WIS] 2.8 c G56 [W13] 'kg OX/kg fuel c C57 [we ] 'Chamber Pressure c F57 [WIS] 500 c G 57 [W1 3) ' psia c C58 [W8] 'Nozzle Expansion Ratio c F58 [W15] 50 c C 59 [W8] 'Specific Inpulse c F59 [WIS] 356.05 c G59 [ Vr 1 3 ] 'sec c C60 [we ] 'Characteristic Velocity c F 60 (W1 5 ] 6047 c G60 (W1 3] 'ft/sec c A62 [W7] 'Case U c C62 (W8] 'LH2+L0X (out)/EARTH TRANSP. CH4+ISRU LOX (in) c E64 [Wll] "[Refrig.] c A65 [W7 ] ' Fuel (Outbound) c D65 t W 8] "H2 c E65 [Wll] •Yes c A66 [W7] Oxidizer(Outbound) c D66 [W 8] "02 c E66 [Hll] " Yes c C67 [W8] 'Oxidizer/Fuel Ratio c F67 [W15] 5.6 c G67 (W1 3] 'kg OX/kg fuel c C68 [WB] 'Chamber Pressure c F68 [WIS] 500 c G68 (W13 'psia c C69 (W8) 'Nozzle Expansion Satlo c F69 [¥15 50 c C70 I we] 'Specific Iapulse c F70 [W15 44)5.95 c G70 [ W13 'sec c C71 [W8J 'Characteristic Velocity c F71 (W15 7555 c G71 (VI 3 'ft/sec c E73 1 w a 1 "[Refrig.) c F73 (W15 "(Surf. Refrig.) c G73 (W1 3 •(ISRU) c A74 [W7] Fuel (Return) c D74 [W8] "CH4 c E7 C 11 [KI1] 'DATABASE RETRIEVAL INFORMATION c J 2 (Wll) 'Field list for paradox file PROPEL.DB converted to c J3 t W 11] ' dBase III foraat in file PROPELDB.DBF c K6 (Wll) ' FUEL c L6 [Wll] 'Character c M6 [Wll] 4 c N6 (Wll) NA c 06 (FA) [Wll] N A c P6 [Wll] NA c K7 [Wll] •OXIDIZER c L7 [Wll] 'Character c M7 [Wll] 5 c N7 [Wll] NA c 07 (FA) [Wll] NA c P7 [Wll] NA c K 8 (S3) [Wll] 'OFRATIO c L8 [Wll] 'Numeric c M8: (Wll) '19, c P23 (S3) (Wll] 'CSTAR c Q23 (S3) (Wll] 'ISP c R23 (Wll] 'ISPVAC c L24 (Wll] 500 c N24 [Wll] SO c m [Wllj 'Output Range c 143 [Wll] 'FUEL c J43 [Wll] 'OXIDIZER c KO (S3) (Wll) 'OFRATIO c L43 (S3) [Wll] 'PC c M43 (S3) [Wll] 'TC c N43 [Wll] 'ARATIO c 043 (S3) (Wll] 'PRATIO c P43 (S3) [Wll] 'CSTAR c Q43 (S3) [Wll] 'ISP c R43 [Wll] 'ISPVAC c It* (L) [Wll] 'CH30H c (L) [Wll] 'N204 c K44 (G) [Wll] 3.1266 c L44 (G) [Wll] 500 c M44 (G) [Wll] 3183.2 c N44 (G) [Wll] 50 c 0 D C17 1*2] -+ D D17 •02 D E17 •20% A1 D H17 (F3) 6.570 D B18 [*13] 'C6U802(OH)2 D C18 [«2] "* D Die •02 D E18 'Nvlon D H18 (F3) 1.908 D A20 1*3] ' Mars region D B21 (W13] " N2H4 D C21 1*2] "• D D21 '02 D H21 (F3) 1 D B22 (W13] "NH3 D C22 [W2] *+ D D22 •02 D H22 (F3) 1.627 D B23 [*13] "CH'i D C23 1*2] "+ D D23 '02 D H23 (F 3 ) 1 D B2*. [*13] "CO D C24 [W2] •+ D D24 '02 D E2<. 'Carbon monoxide D H24 (F3 ) 1.143 D B25 [*13] "CO D C25 |*2] •+ D D25 ' H202 D H25 (F3) 1.213 D B26 (W13) "CO D C26 [*2) *+ D D26 ' N204 D H26 (F3) 0.821 D B27 (*13] "C2H50H D C27 1*2] "+ D D27 •02 D E 27 'Ethyl alcohol D H27 (F3) 2.087 D B20 [*13] "CH30H D C28 1*2] '+ D D28 •02 D £28 'Methyl alcohol D H28 (F 3) 1.5 D B29 1*13] "CH30H D C29 1*2] "+ D D29 ' H202 D H29 (F3) 3.188 D B30 1*13] "CH4 D C30 1*2] *+ D D30 'H202 D H 30 (F3) 8.5 D B31 1*13] "CH D D31 'N204 D H31 (F3) 5.75 D B32 [W13] "NH3 D C32 [M2] "+ D D32 'H202 D H32 (F3) 4.313 D B33 [H13] "N2H4 D C33 {W2] "+ D D33 'H202 D H33 (F3) 4.25 D B34 [M13] "CH30H D C34 [W2] *+ D D34 'N204 D H34 CF3) 10.422 D B35 (W13) "C2H50H D C35 [W2] -+ D D35 ' N204 D H35 (F3) 5.042 D B36 [W13] "C2H50H D C 36 [W2] •+ D D36 ' H202 D H36 (F3) 5.131 D A39 |W3] 'DATABASE PARAMETERS D A40 (W3] \- D B40 (*13) \- D C40 1*2) \- D D40 \- D E40 \- D F40 \- D G40 \- D H40 \- D A41 (Vi3) ' Chamber Pressures (psla) D B42 (*13) 'Eartb region D E42 '50, 100, 500, 1000, 2000 D B4 3 (Ml3] 'Moon region D £<<3 '0.1, 10, 50, 100 D B44 IW13] 'Mars region D E44 '1, 10, 50, 100, 500 D A46 (W3) ' Nozzle Area Ratios D B47 [Wl3] 'All regions D E<>7 '5, 10, 25, 50, 100, 250 D A49 [W3] ' Oxidizer/Fuel Ratios D B50 (W13) 'All regions use stoichiometric values given above vitb D B51 {W13] 'multipliers of: D ESI '0.3, 0.7, 1.0, 1.5, 2.5, 3.5 D A54 (W3] •PROPELLANT SPECIFIC VOLUME (at a given tank pressure) D ASS (*3) \- D B55 (*13) \" D C55 1*2) \- D D55 \- D E55 \- D F55 \" 0 C55 \- D H55 \- D B56 lW13] '(Saturated liquid values at varying pressures) 231 B57 [K13] '(Represents worst case values in calculations) F59 •Pc (psl) H59 'Ptank (MPa) F60 \- H60 \- D61 (W13] 'Note: Assuae tank F61 (F2) 1 H61 (F6) 1.5*F61"0.101323/14.696 C62 [W2] 'pressure is F62 (F2) 10 H62 (F6) 1.5'F62*0.101325/14.696 C63 [W2] '1.3 tines F63 (F2) 50/1.5 H63 (F6) 50*0.101325/14.696 C64 tW2] 'greater than F64 (F2) 50 H64 (F6) 1. 5"F64*0. 101325/14.696 C65 (W2] 'the chamber F65 (F2) 100 H65 (F6) 1.5*F65"0.101325/14.696 C66 [W2] 'pressure. F66 (F2) 500 H66 (F6) 1.5*F66"0.101325/14.696 A68 [W3J 'Table of Specific Volu»es (m"3/kg) A69 (W3) \= B69 (W13] C69 [W2] \* D69 \ = E69 \ = E70 Tank Pressures in MPa D71 (F6) 0.0103420998911268372 E71 (F6) 0,103420998911266372 F71 (F6) 0.344736663037561241 G71 (F6) 0.517104994556341862 H71 (F6) 1.03420998911268372 171 (F6) 5.17104994556341862 A72 (W3) 'Chemical Species D72 E72 F72 G72 H72 172 B73 {W13 J H2 D73 (F6) 0 ,013076 E73 (F6) 014142 F73 (F6) 015592 G73 (F6) ,016552 H73 (F6) ,022494 173 (F6) ,031888 D74 IW13) 'N2H4 E74 (F6) 0 ,000989 B75 [W13] 'NH3 D75 (F6) 0 ,001377 E75 (F6) 0 ,001468 232 D F75 (F6) 0.001547 D G75 (F6) 0.001583 D H75 (F6) 0.001662 D 175 (F6) 0.002076 D B76 [W13] -02 D D76 (F6) 0.000816 D E76 (F6) 0.000878 D F76 (F6) 0.000934 D G76 (F6) 0.000962 D H76 (F6) 0.001028 D 176 (F6) 0.002293 D B77 {W13} "N204 D D77 0.000687 D E77 (F6) 0.00069 D F77 0.000749 D G77 0.00078 D H77 0.000812 D 177 0.001061 D B78 (W13) "H202 D E78 (F6) 0.000691 D B79 [HI 3] "CH4 D D79 (F6) 0.002215 D E79 (F6) 0.002368 D F79 (F6) 0.002524 D G79 (F6) 0.002601 D H79 (F6) 0.00279 D 179 (F6) 0.006233 D B80 [V13] "SiH4 D B81 (W13] *A1 D B82 [W13) "H2 + Al D B83 [W13] *H2 + Al D B84 [W13] "C6HB02(OH)2 D B85 (W13] "CO D D85 0.001182 D E85 0.001268 D F85 0.00136't D G85 0.001417 D H85 0.001541 D 185 0.003322 D B86 [W13] -C2H50H D D86 (F6) 0.00125 D E86 (F6) 0.001323 V F86 (F6) 0.001392 D 686 (F6) 0.001427 D H86 (F6) 0.001512 D 186 (F6) 0.002126 D B87 (W13] -CH30H D D87 0.001263 D E87 0.001332 D FB7 0.001395 D G87 0.001432 D H87 0.001502 D 187 0.00197 D A89 [W3] 'Table of Sto D A90 [W3] 233 B90 1*13] C90 [W2) > D90 \» E90 \» E91 Tank Pressures in MPa D92 (F6) 0.0103420998911266372 £92 (F6) 0.103420998911268372 F92 (F6) 0.344736663037561241 692 (F6) 0.517104994356341862 H92 (F6) 1.03420996911268372 192 (F6) 5.17104994556341862 A93 (W3) 'Cheaic&l Species D93 E93 F93 G93 H93 193 B94 [HI3] *H2 D94 (F2) 14.44 E94 (F2) 20.35 F94 (F2) 25.22 G94 (F2) 27.29 H94 (F2) 31.44 194 (F2) 32.94 B95 [W13] -N2B4 B96 {W13] *NH3 D96 (F2) 202.27 E96 (F2) 240.2 F96 ( F2) 267,4 G96 (F2) 278.21 H96 (F2) 299.14 196 (F2) 363.64 B97 [W13) "02 D97 (F2) 72.47 E97 (F 2) 90.35 F97 (F2) 103.64 G97 (F 2) 109.2 H97 (F 2) 120.19 197 (F2) 154.58 B98 [W13] -N204 D98 (F2) 263.91 E98 (F2) 291.69 F98 (F2) 322.25 G98 (F2) 333.36 H98 (F2) 347.25 198 (F2) 405.59 B99 [W13] -H202 B100 : [W13) "CH4 DlOO : (F2) 90.68 ElOO : (F2) 111.66 FlOO : (F2) 128.82 GlOO : (F2) 135.85 HlOO : (F2) 149.84 IlOO : (F2) 190.55 TT C**i CN v-/«H rH W H ^ u« - a CM - Q M « V •-« w Q; IO (A4 O [i| • V)M •COH mCQQ) r> A& • V) - - CO - E B24 [W18] 'MASS MARGIN E C24 (F2) [WIS] 0.05*@SUM(CI 2..C23) E B26 [W18) •SUBTOTALS E C26 (F2) [W13] ?SUM(C12..C24) E D26 (F2) [W16) £SUH(D12..024) E E26 (F2) (WIS] @SUM(E12..E24) E B28 (WIS) •HEAT SOURCE E C28 (F2) (W13] 8.4"E26/1000 E B29 ["18) •POWER SOURCE E C29 E C47 (F2) (Wl3) 0.226*D3B E B48 [¥18] 'MEMBRANE SEPARATOR E C46 [HI 3] 1.36*D38/10 E B49 [WIS] 'RECYCLE COMPRESSOR E C49 (F2) [W13) 1,0 E C73 (F2) [W13] "Mass (kg) E D73 (F2) [W16] "Watts-electric E E73 (F2> [WIS] "Watts-theraal E B74 C WIS] •FILTER E C74 (F2) IW13] +D69*0.0433 E D74 (F2) [W16] 16.03*(C74) " 1.3 E B75 [WIS) 'C02 ACCUMULATOR E C75 (F2) [W13] 12.3*D69/10 E D75 (F2) [W16] 7.65*(C75)" 1.5 E B76 (WIS] 'HEAT EXCHANGER E C76 (F2) [W13] +D69*0.32*1.476 E E76 (F2) (W15] 61.9*D69*0.327 E B77 [WIS] 'ZIRCONIA CELLS E C77 (F2) [W13] 17.5*069/10 E D77 (F2) [W16] 11,61*(C77)-1.5 E E77 (F2) [W15] 3206*D69/10 E B70 [WIS] 'RADIATOR E C78 (F2) [W13] 0.226*D69 E B79 [WIS] 'MEMBRANE SEPARATOR E C79 [W13] 1.36*069/10 E B80 [WIS] •RECYCLE COMPRESSOR E C80 (F2) [W13] 1.045* D69+10 E D80 (F2) [W16] 2.16*(C80)"1.3 E B81 [WIS] •02 COMPRESSOR E C81 ( F2) [W13] 1.65*D69+13 E E81 (F2) [W15] 945.75*D69/10 E BB2 (WIS] •INT. STORAGE E C82 (F2) [W13] 5.36 E B83 [WIS] •PIPING / TUBING E C83 (F2) [W13] 0.6*D69 E B84 [WIS] •VALVES E D84 (F2) [W16] 31.6 E B85 [WIS] 'COMPUTER + COMM.' E C85 (F2) [W13] 0.1«0SUM(C74.,C83) E B86 [WIS] •MASS MARGIN E C66 (F2) [W13] 0.05"£SUM(C74.,C85) E B88 [WIS] 'SUBTOTALS E C88 (F2) [W13] §SUM(C74..C86) E D88 (F2) [W16] @SUM{074..D86) E EBB ( F2) [W13] fISUM(E74. .E86) E B90 [WIS] 'HEAT SOURCE E C90 (F2) [W13] 8.A"E8B/1000 E B91 [WIS] •POWER SOURCE E C91 (F2) [W13] +D71"D8B/1000 E A93 1W5] 'TOTAL MASS E C93 (FA) [W13] §SUM(C8B.,C91)*100/197 E D93 (F4) [W16] E E93 (FA) [WIS] 'CASE 3 E A94 [W5] \* E B94 [WIS] \» E C94 [ W13] E D94 IW16] E E94 [ W15] \" E A97 [ W 5 J \- E B97 [WIS] \- 242 E C97: ["13) \" E A98: [W5J ' CASE 4 MASSES E E98: ["15] "REQUIRED E A99: IW5] \- E B99: [WIS) \- E C99: ["13] \- E E99: ("131 "PROD. RATE E B100 ["18] 'PRODUCTION RATE E ClOO ["13] '(kg 02/day) E DlOO ["16) 10 E ElOO (Fl) ("15] +SL:SCS119/D101 E BIO 1 ["18] 'PRODUCTION TIME E C101 ("13] '(days) E D101 l"16) 200 E B102 (F2) ("18] 'POWER SUPPLY SP. WT. (kg/kwe) E D102 ("16] 30 E A104 ("5] "FACTORY COMPONENTS E C104 (F2) [H13] "Mass (kg) E DIOA (F2) ("16] ""atts-electric E E104 (F2) ["15] "Watts-thermal E B105 1*18] 'FILTER E C105 (F2) t W13] +D100*0.0455 E D105 (F2) (W16] 16.03"(C105)*1.5 E B106 l"18] 'C02 ACCUMULATOR E C106 (F2) ["13] 12.3* DlOO/10 E D106 (F2) ["16) 7.65"(C106)"1,5 E B107 ("18] •HEAT EXCHANGER E C107 (F2) [W13 ] + D100-0.52M.476 E E107 (F2) ["15] 61.9-D100-0.527 E BIOS ("181 'ZIRCONIA CELLS E C108 (F2) ("13] 17,5*D100/10 E D108 (F2) ["16] 11.61"(C108)-1.5 E El 08 (F2) ["15] 3206-D100/10 E B109 1" 18 ] 'RADIATOR E C109 (F2) ("13] 0.226'DIOO E B110 ("18| 'MEMBRANE SEPARATOR E Clio ("13] 1.36"D100/10 E Bill ["18] •RECYCLE COMPRESSOR E cm (F2) ("13] 1.045"D100+10 E Dill (F2) ["16] 2.16'(C111)*1.5 E B112 ["18) '02 COMPRESSOR E CI 12 (F2) ["13] 1.65"D100+15 E El 1 2 (F2) ("15] 945.75"D100/10 E B113 f"18] 'INT. STORAGE E CI 1 3 (F2) ["13] 5.36 E B114 ("18] •PIPING / TUBING E C114 (F2) ["13] 0.6*D100 E B115 [W18] 'VALVES E D115 (F2) ["16] 31.6 E D116 ["18] •COMPUTER + COMM. E CI 16 (F2 ) ["13] 0.1'$SUM(C105..C114) E B117 ["18] 'MASS MARGIN E CI 17 (F2) [W13) 0.05*$SUM(C105..CI 16) E B119 ("18] 'SUBTOTALS E C119 (F2) ("13] fISUM(C105 . . CI 17) 243 E D119 (F2) [W16] £SUM(D103..D117) E E119 (F2) (WIS] $SUM(E105,,E117) E B121 [WIS] 'HEAT SOURCE E C121 (F2) (W13] 8.4«En9/1000 E B122 [f 18] •POWER SOURCE E C122 (F2) (W13] +D102*D119/1000 E A124 [W5] 'TOTAL MASS E C124 (F<0 [W13] §SUM(C119..C122) E D124 (FA) [W16] " > E E124 (F4) [WIS] 'CASE U E A123 [W5] \" E B125 [WIS] E C125 [W13] \» E D12S (W16] \» E El 25 [W15] A1 (¥51 'SHEET U6 - SUPPORT COMPONENTS A4 [W5) 'REFRIGERATION REQUIREMENTS FOR PROPELLANT WHILE ON THE MARTIAN A5 [W5] ' SURFACE (BASED ON RADIATIVE AND FREE-CONVECTIVE HEAT FLUX) B7 [W13J 'ATMOSPHERIC DENSITY [kg/«3J E7 (F7) [WLL] 6000/(189*E8) B8 [W13) •AMBIENT T (PEAK) [K) E8 (F2) [Wll] 260 B9 [W13] 'VISCOSITY [kg/as] E9 (F7) [Wll) 1.31E-05 BIO [W13] 'ISOBARIC COMP [1/K) ElO (F7) [Wll] 1/E8 B11 [WL 3] 'PRANDTL NUMBER Ell (F5) (W11J 0.764 B12 [ Wl 3] 'THERMAL CONDUCTIVITY IW/BK] F El 2 [Wll] 0.01385 F A1A [W5J ' DIAMETERS OF TANKS (FOUR PER STAGE) F D13 (Wll) • Case 1- F E15 [Wll] ' F F15 (Will • Case 2- F G15 [Wll] F D16 [Will "OXIDIZER F E16 [Wll] "FUEL F F16 [Will -OXIDIZER F G16 J W1 1 ] " FUEL F B17 (F 2) [W13) •DLEG6IV [in] F D17 F D23 (S2) [Wll] (6«$I:$K$78/($G:$CS19»ePI))" 1/3) F E23 (S2) [Wll] (6*$I:SKS77/($G: SCS X7-#PI))- 1/3) F F23 (S2) (Wll] (6*$J:SKS78/($G:SDS19*@PI))* 1/3) F G23 (S2) [Wll] (6"$J:$K$77/($G:$D$17*gPI))" 1/3) F B24 (F2) [W13] 1DLEG5I (•] F D24 (S2) [Wll] (6*SI:SJS7B/(SG:SCS19*@PI))* 1/3) F E24 (S2) [Wll] (6*$I:SJ$77/(SG:SCS17"£PI))" 1/3) F F24 (S2) [Wll] (6*$J:SJS78/C SG:$DS19"?PI))- 1/3) F G24 (S2) [Wll] (6*$J:$J$77/(SG:SDS17*£PI))* 1/3) F B25 [W13] 'DLEG4IV (•] F D25 (S2) [Wll] (6* SI:STS78/C SG:SCS 19*flPI))* 1/3) F E25 (S2) [Wll] (6* SI:STS77/(SG: SCS17"pPI))* 1/3) F F25 (S2) [Wll] (6*SJ:STS78/(SG:SDS19*£PI))* 1/3) F G25 CS2) [Wll] (6* SJ:STS77/(SG:SDS17'?PI))- 1/3) F B26 (F2) [W13] 'DLEG4II1 [•] F D26 (S2) (Wll] (6"SI:SSS78/C SG:SCS19*§PI))- 1/3) F E26 (S2) [Wll] (6* SI:SSS77/(SG:SCS17'gPI))* 1/3) F F26 (S2) [Wll] (6'SJ:SSS78/[SG:SDS19*(1PI))" 1/3) F G26 (S2) [Wll] (6"SJ:SSS77/(SG:SDS17"@PI))" 1/3) F B27 (F2) [W1 3] 'DLEG4II [n] F D27 (S2) [Wll] (6-SI:SRS78/(SG:SCS19"PPI))* 1/3) F E27 (S 2) [Wll] (6 • S I:SRS77/(SG: SCS17"(1PI))* 1/3) F F27 C S 2) [Wll] C 6•SJ:SRS78/[SG: SDS19"gPI))" 1/3) F G27 (S 2) [Will (6*SJ:SRS77/(SG: SDS17*pPI))* 1/3) F B28 (F2) 1W1 3] 'DLEG4I [n] F D28 (S2) [Wll] (6-SI:SQS78/(SG:SCS 19 "(iPI))_ 1/3) F E28 (S2) [Wll] (6• SI: SQS77/(SG: SCS17"gPI)) * 1/3) F F26 (S2) [Will (6* SJ:SQS78/(SG:SDS19*£PI))" 1/3) F G28 (S2) [Wll] (6•SJ:SQS77/(SG:SDS17'gPI))- 1/3) F A30 [W5] ' GRASHOFF NUMBERS F D31 [Wll] • Case 1- F E 31 twin • F F31 [Wll] • Case 2- F G31 (Will • F D32 [Wll] "OXIDIZER F E32 [Wll] •FUEL F F32 (Wll] 'OXIDIZER F G32 [Wll] "FUEL F B33 (F2) [W13] 1 GR6IV F D33 (S2) [Wll] (SES7/SES9)*2,SB:SFS12*SES10*(SES8 •SG:SCS63)"D17"3 F E33 (S 2) [Wll] (SES7/SES9)* 2 * SB :SFS12* SES10*(SES8 - S G:SCS63)"E17 * 3 F F33 (S 2) [Wll] (SES7/SES9K 2*SB:SFS12 *SES10*(SES6 •SG: SDS65)"F17*3 F G33 (S2) [Wll) (SES7/SES9)•2,SB:SFS12*SES10*(SES8 •SG: SDS63)*G17 *3 F B34 [W13] 'GR6III F D34 (S2) [Wll] (SES7/SES9)-2,SB:SFS12*SES10*(SES8 •SG:SCS65)*D10 *3 F E34 < S2) [Wll] (SES7/SES9)"2* SB:SFS12*SES10* (SES8 •$G:SCS63) "E18* 3 F F34 (S 2) [Wll] (SES7/SES9)'2*SB: SFS12 aSES10*(SES8 •SG:SDS65)"F18*3 F G34 (S2) [Wll] (SES7/$E$9)*2*SB:SFS12*SES10*(SES8 •SG :SDS63) "G18" 3 F B35 (W13] 'GR6I1 F D35 < S 2) [Wll] (SES7/SES9)"2"SB:SFS12*SES10'(SES8 •$ G:SCS65)*D19* 3 F E35 (S 2) [Wll] (SES7/SES9)*2"SB:SFS12* SES10*(SES6 •SG:SCS63)•E19" 3 F F35 (S 2) [Wll] (SES7/SES9)'2* SB:SFS12*SES10" (SES8 •SG:SDS65)"F19*3 F G35 (S2) [Wll] (SE$7/SE$9)"2*SB:SFS12,SES10*(SES8 •SG: SDS63)"G19" 3 F B36 IW1 3) 'GR6I F D36 F E36 (S2) [Wll) (SES7/SES9)"2*SB SFS12•SES10 (SES8-SG SCS63) E20-3 F F36 (S2) [Wll] (SES7/SES9)"2 "SB SFS12•SESIO (SES8-$G:SDS63) F20-3 F G36 (S2) [Wll] (SES7/SES9)-2 •SB:$F$12*SES10 (SES8-$G:SDS63) G20-3 F B37 [W13J 'GR3IV F D37 (S2) [Wll] (SES7/SES9)"2 •SB SFS12•SESIO (SES8-SG SCS65) D21 *3 F E37 (S2) [Wll] (SES7/SES9)-2'SB SFS12•SESIO (SES8-$G:SCS63) E21 -3 F F37 (S2) [Wll] (SES7/SES9)-2 •SB SFS12•SESIO (SES8-SG SDS63) F21 -3 F G37 (S2) [Wll] (SES7/SES9)-2 • SB SFS12•SESIO (SES8-SG SDS63) G21 -3 F B38 IW13] •GR5III F D38 (S2) [Wll] (SES7/SES9)*2 •SB SFS12•SESIO (SES8-SG:SCS65) D22*3 F E30 (S2) [Wll] (SES7/SES9)-2 "SB SFS12•SESIO (SES8-SG:$CS63) E22-3 F F38 (S2) [Wll] (SES7/SE59)-2 •SB SFS12•SESIO (SES8-SG SDS65) F22-3 F G38 (S2) [Wll] (SES7/SES9)"2 •SB SFS12•SESIO (SES8-SG:SDS63) G22-3 F B39 [W13] 'GR5II F D39 (S2) [Wll] (SES7/SES9)*2 • SB SFS12•SESIO (SES8-SG:SCS65) D23"3 F E39 CS2) [Wll] (SES7/SES9)-2 •SB SFS12•SESIO (SES8-SG SCS63) E23-3 F F39 (S 2) [Wll] (SES7/SES9)•2 •SB SFS12•SESIO (SES8-SG SDS65) F23 -3 F G39 CS2) [Wll] (SES7/SES9)*2 •SB SFS12•SESIO (SES8-SG SDS63) G23-3 F B40 [W1 3] •GR5I F D40 (S2) [Wll] (SES7/SES9)•2 •SB SFS12•SESIO (SES8-SG SCS65) D24 -3 F E40 (S2) [Wll] (SES7/SES9)"2 •SB SFS12•SESIO (SES8-SG SCS63) E24 A3 F FAO (S2) [Wll] (SES7/SES9)"2 •SB SFS12•SESIO (SES8-SG SDS65) F24 -3 F G40 (S2) [Wll] (SES7/SES9)"2 •SB SFS12•SESIO (SES8-SG SDS63) G24"3 F B A1 t W13] 1 GR41V F D4 1 C S 2) [Wll] (SES7/SES9)-2 •SB SFS12•SESIO (SES8-SG SCS65) D25"3 F EM CS2) [Wll] (SES7/SES9)"2 •SB SFS12•SESIO (SES8-SG SCS63) E25-3 F F41 (S2) [Wll] (SES7/SES9)*2 •SB SFS12•SESIO (SES8-SG SDS65) F25 *3 F G41 (S2) [WU] (SES7/SES9)"2 •SB SFS12•SESIO (SES8-SG SDS63) G25 *3 F B42 (W13] 1 GR4111 F D42 C S 2) [Wll] (SES7/SES9)•2 •SB SFS12•SESIO (SES8-SG SCS65) D26 -3 F E42 CS2) [Wll] (SES7/SES9)*2 •SB SFS12•SESIO (SES8-SG SCS63) E26 -3 F F42 C S 2) [Wll] (SES7/SES9)*2 •SB SFS12•SESIO (SES8-SG SDS65) F26 -3 F G42 (S2) [Wll] (SES7/SES9)*2 •SB SFS12•SESIO (SES8-SG:SDS63) G26*3 F BO [W13 ] 'GRA11 F D4 3 (S 2) [Wll] (SES7/SES9)"2 •SB SFS12•SESIO (SES8-SG:SCS65) D27"3 F EA3 (S2) [Wll] (SES7/SES9)•2 •SB SFS12•SESIO (SES8-SG SCS63) E27 -3 F F43 t S2) [Wll] (SES7/SES9)*2 •SB SFS12•SESIO (SES8-SG SDS65) F27 -3 F Gt3 C S2) [Wll] (SES7/SES9)*2 •SB SFS12•SESIO (SES8-SG:SDS63) G27*3 F B4 4 (W13) 1 G R 41 F X)hU (S2) [Wll] (SES7/SES9)*2 •SB SFS12•SESIO (SES8-SG SCS63) D28 -3 F E44 (S2) [Wll] (SES7/SES9)-2 •SB SFS12•SESIO (SES8-SG SCS63) E28-3 F F44 t S 2) [Wll] (SES7/SES9)•2 •SB SFS12•SESIO (SES8-SG:SDS65) F28 -3 F G44 (S2) [Wll] (SES7/SES9)*2 •SB SFS12•SESIO (SES8-SG:SDS63) G28"3 F A46 [W5] • RAYLEIGH NUMBERS F D47 [Wll] ' Case 1- F E47 [W11J > F FA7 [Wll] • ---Case 2- F G47 [Wll] > F D48 [Wll] •OXIDIZER F E^t 8 [Wll] -FUEL F F48 [Wll] -OXIDIZER F G48 [Wll] -FUEL F B49 [W13] •RA6IV F 9 (S2) [Wll] +D33"SES11 F EA9 (S2) [Wll] +E33*SES11 247 F F49 CS2) [Mil +F33 SES11 F G49 (S2) [Mil +G33 SESll F B50 [*13] •RA6I11 F D50 (S2) [Wll +D34 SESll F E50 (S2) [Wll +E34 SESll F F50 (S2) [Wll +F34 SESll F GSO (S2) [Nil •G3 [Wll + E39 SESll F F55 (S2) (Wll + F39 SESll F G55 (S2) [Wll + G39 SESll F B56 (W131 •RA I F D56 (S 2) [Wll + D40 SESll F E56 (S 2) [W11 + E40 SESll F F56 (S 2) [Wll + F40 SESll F G56 (S 2) [Wll + G40 SESll F B57 t«13) •RA IV F D57 (S2) [Wll + D4 1 SESll F E57 (S2) [Wll + E41 SESll F F57 (S2) [Wll +F41 SESll F G57 (S2) [Wll +G<<1 SESll F B58 t W13] 'RA III F D58 (S2) [Wll +D42 SESll F E58 (S2) (Wll •Ei.2 SESll F F58 (S 2) [Wll +FA 2 SESll F G58 (S2) [Wll •G42 SESll F B59 [W13] 'RA II F D59 (S2) [Wll +D43 SESll F E59 (S2) [Wll +E43 SESll F F59 (S 2) [Wll +F^ 3 SESll F G59 (S2) [Wll +C.3 SESll F B60 t W13 ] 'RA I F D60 (S2) [Wll +D^i ^ SESll F E50 (S2) [W11 +E44 SESll 248 F F60 CS2) (Wll] +F44"SES 11 F G60 (S2) [Wll] +G44»SESH F A62 [W5] • NUSSELT NUMBERS F D63 [Wll] ' ---Case 1- F E63 [Wll] • F F63 [Wll] ' -—Case 2- F G63 [Wll] ' F D64 [Wll] "OXIDIZER F E64 [Wll] •FUEL F F64 [Wll] -OXIDIZER F G64 [Wll] •FUEL F B65 [W13] 'NU6IV F D65 (S2) [Wll] 2+(0.509"D49 "0.25/(1 (0.469/SES11 "(9/16) -(4/9) F E65 (S2) [Wll] 2+(0.589"E49 *0.25/(1 (0.469/SES11 •(9/16) "(479) F F65 F G73 (S2) [Wll] 2+(0.589*G57"0.25/(l+(0,469/SE$U)-(9/16))-(4/9)) F B74 [*13] 1 NU^t III F D74 (S2) [Wll] 2+(0.589,D58"0.25/(l+(0.469/SES11)' (9/16)) (4/9)) F E74 (S2) [Wll] 2+(O.589*E50-O.25/(l+(O,469/SES11)' (9/16)) (A/9)) F F74 (S2) [Wll] 2+(0.589"F58"0.25/(l+(0. 469/SES 11)'(9/16)) (4/9)) F G74 (S2) [Wll] 2+(0.589*G58-0.25/(l+(0,469/SESll)' (9/16)) (4/9)) F B75 [W13] 'NU4II F D75 (S2) [Wll] 2+(0.589*DS9"0.23/(l+(0.469/SES11) ' (9/16)) (4/9)) F E75 (S2) [Wll] 2+(0.589*E59*0.25/(l+(0,469/SES 11)'(9/16)) (4/9)) F F75 (S2) [Wll] 2+(0.589*F59-0.25/(l+(0.469/SES11)' (9/16)) (4/9)) F G75 (S2) [Wll] 2+(0.589"G39*0.25/U +(0. 469/SES11)' (9/16)) (4/9)) F B76 [W13] ' NU4I F D76 (S2) [Wll] 2+(0.389*D60"0.25/(l+(0.469/SESll)* (9/16)) (4/9)) F E76 (S2) [Wll] 2+(0.589"E60-0.25/(l+<0.469/SES U)'(9/16)) (4/9)) F F76 (S2) (Wll) 2+(0.589*F60*0.25/(l+(0,469/SES11)' (9/16)) (4/9)) F G76 (S2) [Wll] 2+(0.S89*G60"0.25/(l+(0,469/SES11)' (9/16)) (4/9)) F A7 8 [ W5] ' CONVECTION COEFFICIENTS F D79 [WU) ' Case 1 - F E79 [WU] F F79 [WU] ' Case 2- F G79 (WU] F D80 [WU] 'OXIDIZER F E80 [WU] "FUEL F F80 [WU] 'OXIDIZER F G80 [WU] "FUEL F B81 [ W13] 'H61V [W/n2K] F D81 (S2) (Wll) 0IF(D17=O,0I+D65"$E$12/D17) F E81 (S2) [Wll] @IF(E17«0,0,+E65*SES12/E17) F FBI (S2) (Wll) 0IF(F17«O,0.+F65-SES12/F17) F G01 (S2) [Wll] eiF(G17=0,0,+G65*5ES12/G17) F B82 [ W13] • H6111 [W/n2K] F D82 (S 2) (Wll) PIF(D18*0,0,+D66"SES12/D18) F E62 F B114 [f 13] •MR0X6III [kg] F D114 < S 2) [WUJ +D98'$G:SCS«VSG:SC$52 F El 14 (S2 ) [Nil] •E98*$G:$C$'i3/$G:$C$51 F FU4 (S2) [Nil] +F98*$G:$D$44/$G:$DS52 F G114 G B17 (W18) 'DF(return) [kg/»3] G C17 (F3) [Nil] 1/SD:SIS 112 G D17 (F3) (Nil) 1/SD:SIS 117 G E17 (F3) [Nil] 1/SD:$IS 122 G F17 C F3) [Will 1/SD:SIS 127 G BIS [WIS] 'DOX(outbd) [kg/«3] G C18 (F3) [Wll] 1/SD:$I$113 G D18 (F3) [Wll] 1/SD:SIS118 G El 8 (F3) [Wll] 1/$D:SIS 123 G F18 (F3) [Wll] 1/SD:$IS 128 G B19 [ W18) 'DOX(retrn) (kg/«3) G C19 (F3) [Wll] 1/SD:SIS113 G D19 (F3) [Wll] 1/SD:SIS11S G E19 (F3) (Wll] 1/SD:SIS123 G F19 (F3) [Wll] 1/SD:SIS128 G B20 ( W 1 8] 'EMISSF G C20 (F3) (Wll] 0.04 G D20 (F3) [Wll] 0.04 G E20 (F3) (Wll] 0.04 G F20 (F3) (Wll] 0.04 G B21 1*181 'EMISSOX G C21 (F3) (Wll] 0,04 G D21 (F3) [Wll] 0.04 G E21 (F3) [Wll] 0.04 G F21 (F3) [Wll] 0.04 G B22 [WIS] 'EC G C22 (F3) (Wll) +C66"0.0177 G D22 (F3) (Wll] +D66*0.0177 G E22 (F3) [Wll] +E66*0.0177 G F22 (F 3) (Wll] +F66-0.0177 G B23 [ W18] 'FAF G C23 (F3) [Wll] 0.5 G D23 (F3) [Wll] 0.5 G E23 (F3) [Wll] 0.5 G F23 (F3) [Wll] 0.5 G B24 [ W1 81 'FAOX G C24 (F3) [Wll] 0.5 G D24 (F3) [Wll] 0.5 G E24 G B28 [W1B] 'FOXL(return) G C28 (F3) HI 1] 0.02 G D28 (F3) HU] 0.02 G E2B (F3) Nil] 0.02 G F28 (F3) HU] 0.02 G B29 [HIS] 'FSC(press.) G C29 (F3) Nil] 0.025 G D29 (F3) Hll] 0.025 G E29 (F3) HI 1] 0.025 G F29 CF3) Hll] 0.025 G B30 [H16] 'FSC(puap) G C30 (F3) Hll] 0.057 G D30 (F3) Hll] 0.057 G E30 (F3) Hll] 0.057 G F30 (F3) Hll] 0.057 G B31 [H18] ' FSF G C31 (F3) Hll] 1 G D31 (F3) Hll] 1 G E31 (F3) Hll) 1 G F31 (F3) Hll] 1 G B32 [W18] ' FSOX G C 3 2 (F3) Hll] 1 G D32 (F3) Hll] 1 G E32 (F3) Hll] 1 G F32 (F3) Hll] 1 G B33 1H18] 'FTC G C 3 3 (F3) Hll] 1 G D33 (F3) Hll] 1 G E33 CF3) HI I] 1 G F 3 3 (F3) Hll] 1 G B3 G:B49 (WIS) 'NFT G:C49 [Wll] 4 G: D49 [Wll] 4 G:E49 [Wll) 4 G:F49 [Wll] 4 G:B50 [W18] ' NOXT G: CSO [Wll] 4 G:D50 [Wll] 4 G: ESO [Wll] t< G: F50 [Wll] 4 G: B51 [WIS] 1 NRF G:C51 (F3) [Wll] 0.225;Efficiency is chosen so combined refrigeration mass is a pprox 2000 kg D51 (F3) [Wll] 0.225 E51 (F3) [Wll] 0 ,225 F51 (F3) [Wll] 0 ,225 B52 (WIS] 'NROX C52 (F3) (Wll] 0 225 D52 (F3) [Wll] 0 225 E52 (F 3) [Wll) 0 225 F52 (F3) [Wll] 0 225 B53 (W]8) 'OXTC C53 (F3) [Wll] 1 D53 (F3) [Wll] 1 E53 (F3) [Wll) 1 F53 (F3) [Wll] 1 B54 [W18] 'OXTR C54 (F3) [Wll] 0 D54 (F3) [Wll) 0 E54 (F3) [Wll] 0 F54 (F3) [Wll] 0 B55 [ W18) 'PF(outbound) C55 (F3) [Wll] 100/(SC:SFS7+1) D55 (F3) [Wll] 100/(SC:SFS27+1) E55 (F3) [Wll] 100/(SC:SFS47+1) F55 (F3) [Wll] 100/(SC:SFS67+ 1) B56 f W18] 'PF(return) C56 (F3) [Wll] 100/(SC: SFS16 + 1) D56 (F3) [Wll] 100/(SC:SFS36+1) E56 (F3) [Wll] 100/(SC: SFS56 + 1) F56 (F3) [Wll] 100/(SC:SFS76+1) B57 [W18) 1PSTOR [Pa] C 57 (S 2) [Wll] +SC:SFS8/0.00014504811 D57 CS2) [Wll] •SCiSFSS/O.00014504811 E57 (S2) [Wll] +SC:SFS8/0.00014504811 F57 (S2) [Wll] +SC:SF$8/0.00014504811 B58 [W16] 'SBC [W/B2K4] C58 (S2) [Wll] 5.669E-08 D58 (S 2) [Wll] 5.669E-08 E58 (S2) [Wll] 5.669E-08 F58 (S2) [Wll] 5.669E-08 B59 [W18] 'SCE C59 (F3) [Wll] 0.026 D59 (F3) [Wll] 0.026 E59 (F3) [Wll] 0.026 F59 (F3) [Wll] 0.026 259 B60 WIS] 'SCS C60 F3) [Wll] 0.032 D60 F3) [Wll] 0.032 £60 F3) [Wll] 0.032 F60 F3) [Wll] 0.032 B61 W18] 'SFE [W/02] C61 F3) [Wll] 1386 D61 F3) [Wll] 1386 E61 F3) [Wll] 1388 F61 F3) [Wll] 1388 B62 WIS] 'TF(outbound) (K) C62 Wll] fINDEX(SD:SBS92..SD:SIS108,SI: SCS33,SD:SFS112)-5 D62 Wll] gINDEX(SD:SBS92..SD:SIS108,SJ! SCS33ISD:SFS117)-5 E62 Wll] @1NDEX(SD:SBS92..SD:SIS108,SK: SCS33.SD:SFS122)-5 F62 Wll] £1NDEX(SD:SBS92..SD:SIS108,SL: SCS33,SD:SFS127)-5 B63 W18] 'TF(return) [K] C63 Will £ IN D E X(SD:SBS92..SD:SIS108,SI: SXS33.SD:SFS114)-5 D63 Wll] glNDEX(SD:SBS92..SD:SI$108,SJ:SXS33ISD:SFS119)-5 E63 Wll] gINDEX(SD:SBS92..SD:SIS108,SK: SXS33,SD:SFS12A)-5 F63 Wll) gINDEX(SD:SBS92.,SD:SIS108,SL: SXS33,SD:SFS129)-5 B64 Wl8] 'TOX(outbound) [K] C64 Wll] (ilNDEX(SD:SBS92..SD:SIS108,SI: SCS33.SD:SFS113)-5 D6 G E73 [Wll 18 G F73 [Ml 1 18 G B75 [WIS ' TITANIUM AL1 G B76 [W18 'FTSFPRESS G C76 [Wll 1 G D76 [Wll 1 G E76 [WU 1 G F76 [Wll 1 G B77 [W18 'FTSFPUMP G C77 [Wll 3 G D77 [Wll 3 G E77 [Wll 3 G F77 [Wll 3 G B78 [W18 'DFT [kg/«3] G C78 [ W11 7803 G D78 [Wll 7803 G E78 [Wll 7803 G F78 [Wll 7603 G B79 [W18 'S1GFT [N/n2] G C79 t S 2) [Wll] 620000000 G D79 (S 2) [Wll] 620000000 G E79 (S2) [Wll] 620000000 G F79 (S 2) [Wll] 620000000 G BB1 [W18 • TITANIUM ALI G B62 [Wis •OXTSFPRESS G C82 I W11 1 G D82 [W1 1 1 G E82 [Wll 1 G F82 [ W1 1 1 G B63 [ W18 'OXTSFPUMP G C83 [Wll 3 G D83 [Wll 3 G E83 [Wll 3 G F83 [Wll 3 G B6« [ W1 8 'DOXT [kg/m3] G C6<( [Wll 7803 G D84 [Wll 7803 G E64 [Wll 7803 G F64 [Wll 7803 G B85 [ W 1 8 •SIGOXT [N/m2] G C85 (S2) [Wll] 620000000 G D85 (S2) [Wll] 620000000 G E85 (S2) (Wll] 620000000 G F85 (S2) [Wll] 620000000 H:A1: (W5] 'SHEET »8 - MODULAR ENGINE DETAILS H:E1: [Mil] ' (Identical for each stage) H:A3: (WSJ 'COMPONENT MASSES H:C3: (W3) '(Based on Pressure-Feed Systea) H:B I B16 [W17 1SFB [W/B2] I C16 [Wll +SG:SCS61*($G:SC$13/C13)"2 I D16 [Wll +SG:SCS61"(SG:$CS13/D15)*2 I E16 [Wll +SG:SCS61»(SG:SCS15/E15)-2 I F16 [Wll +$G:$C$61*(SG:SCS13/F15)-2 I A18 [WS] ' REFRIGERATION MASS PRELIMINARY CALCULATIONS I B19 [W17 ' Oxidizer I E19 [Wll ' Fuel I B20 [W17 'HOX [W/«2] I C20 (F4) (Wll] +$G:SC$58*$G:SC$21"$G:SC$63"4 I E20 [Wll 'HF I F20 (F4) [Wll] +$G:$C$S8,SG:SC$20*$G:$CS63"4 I B21 [W17 'ROXT [B/(kg)*1/3] I C21 (F4) [WU] (0.75/($G:$C$19*$G:$CS50*(iPI))*(1/3) I E21 [Wll ' RFT I F21 (F4) (Wll) (0.75/(SG:SC$17*SG:SC$ 116 [Wl 'SFB [W/n2] J16 [W1 +SG:SCS61"(SG:SCS15/J15)*2 K16 [wi +SG:SCS61,(SG:SCS15/K13)*2 L16 [VII +SG:SCS61"(SG:SCS15/L15)*2 M1 6 [Wl +SG:$C$61"($G:$CS15/M15)"2 HI 8 [W5 ' REFRIGERATION MASS PRELIMINARY CALCULATIONS 119 [W1 ' Oxidizer L19 [Wl • Fuel 120 [Wi 'HOX [W/b2) J20 (FA [Wll] +$G:$C$58*$G:$CS21a$G:$CS65*4 120 [WI •HF M20 (FA [Wll] +SG:$CS58"SG:$CS20»SG:5CS63*A 121 [Wi 'ROXT [m/(kg)* 1/3) J21 (FA [Wll] (0.75/(SG:SCS19*SG:$CS50»pPI))-(1/3) L21 [ Wl 'RFT M21 (FA [Wll] (O.75/($G:$C$17'$G:SC$A9'0PI))-(l/3) 122 (Wl •QIN (W/(kg)"2/3] J22 (FA [Wll] (J16*(1-SG:SC$5A)"SG:$C$2A)"£PI,J21*2 L2 2 [Wl ' QIN M22 (FA [Wll] (M16*(l-SG:SCS3A)"SG:SCS23)*pPI"M21*2 123 (Wl 'QOUT [ " " ] J23 (FA (Wll] (J20,SG:SCS32)"A'gPI°J21'2 L23 [Wl •QOUT M2 3 (FA [Wll] (K20'SG:SCS31)*A*§PI'M21-2 124 [Wl 'QCOOL [ " " ) J2A (FA [Wll] +SG:$CS50*(J22-J23) L2i. [Wl •QCOOL M 2 (FA [Wll] +SG:SCSA9"(H22-M23) 125 (Wl ' MRPRE [kg"1/3] J25 (FA (Wll) (SG:SC$AA"J2A/SG:SC$52) L25 (Wl •MRPRE M25 (FA [Wll) (SG:SCSA3*M2A/SG:5CS51) H27 [W 5 ' ROCKET DESIGN DETAILS (stage dependence) 128 [Wl •FEED SYSTEM USED J28 [Wl "press. K28 (Wl "press. L28 (Wl "press. M28 (Wl "press. 129 (Wl •THRDES [kg] J29 (F0 [Wll] + SG:SC$66*J86 K29 (FO [Wll] +SG:SC$66"K86 1.29 (FO [Wll] •SG:SCS66-L86 M29 (FO [Wll] +SG:$C$66"M86 130 (F2 [Wl 7] 'TIMEB (sec) J30 (F2 [Wll] +J79*J13/J29 K30 (F2 (Wll) eiF(K29»0,0,K79*K13/K29) L30 (F2 (Wll) eiF(L29»0,0,L79'Ll3/L29) M30 (F2 [Wll] §IF(M29»0,0|M79,M13/M29) 131 [Wl ] •OPT. FEED SYSTEM J31 [Wl ] $IF(J30<»(8130.l«J29*-0.75067)," press . punp") K31 (Wl PIF(K29»0," n/a",plF(K30<»(8130.l"K29"-0.75867)," press, putn )) I:L31: [Wl ?IF(L29"0," n/a",piF(L30<"(8130.1'L29--0.75867)," press, pun )) I:M31: [Wl @IF(M29=0," n/a",gIF(M30<-(8130.1"M29--0.75867)," press. pump )) 1:132: (W17] 'PT [MPa] I:J32: (F4) (Wll] 0IF(J28-"PUMP",50*0.101325/14. .696,1.5* $C:SFS17*0. 101325/14.696 ; I:: K32: (F4) (Wll] 0IF(K 28• "PUMP",50*0. 101325/14. .696,1.5* SC:SFS17"0.101325/14.696 ) 1:: L32: (F4) (Wll] 0IF(L28B "PUMP",50*0.101325/14. ,696,1.5* SC:$FS17*0. 101325/14.696 )\ I:: M 32 : (F4) [Wll] 0IF(M28*"PUMP",50*0. 101325/14. .696,1 .5* SC:SFS17*0.101325/14.696 )V I:: 133: (W17] 'PINDEX I;: J33: (FO) [Wll] 0IF(J32> «SD:SIS71,7,0IF(J32>*SD:SHS71, 6,0IF(J32 >« SD:SGS71,5,0I F(J32>=SD:SFS71,4,0IF(J32>»SD:SES71,3,SD:SJS61))))) I:K33: (FO) [Wll] 0IF(K32>«SD:SIS71,7,0IF(K32>»SD:$H$71,6,0IF(K32>»SD:SGS71,5,0I F(K32>=SD:SFS71,4,0IF(K32>«SD:SES71 , 3,SD:SJS61))))) I:L33: (FO) [Wll] 0IF I P16 [W17] 'SFB [W/n2] I Q16 [Wll] +$G:SCS61"(SG:SC$13/Q13)"2 I R16 [Wll] +SG:SCS61"(SG:SCS15/R15)*2 I S16 [Wll] +SG:SC$61,(SG:SC$15/S15)"2 I T16 [Wll] +$G:SCS61,($G:$CS13/T13)"2 I oie [W5] • REFRIGERATION HASS PRELIMINARY CALCULATIONS I P19 [W17] ' Oxidizer I S19 [Wll] ' Fuel I P20 [W17] 'HOX [W/n2] I Q20 (FA) [Wll] +SG:$CS58*$G:SCS21"SG:SCS63"A I S 20 [Wll] •HF I T20 (FA) [Wll] +SG:SC$58"SG:SCS20*SG:SCS63"A I P21 [ W1 7] •ROXT [n/(kg)"1/3] I Q21 (FA) [wii] (o^s/tsGiscs^'SGiScsso-pPDj-d/a) I S 21 [Wll] 'RFT I T21 (FA) [Wll] (0.75/($G:SCS17"$G:$C$A9,flPI))"(l/3) I P22 [ W17] •QIN [W/(kg)'2/3] I Q22 (FA) [Wll] (Q16*(l~SG:SCS5A)"SG:$CS2A)"(iPI,Q21"2 I S22 [Will •QIN I T22 (FA) [Wll] (T16*(1-SG:$CS3A),SG:$CS23)*£PI"T21*2 I P23 [ W17) ' QOUT [ " " ] I Q23 (FA) IW11] (Q20"SGISCS32)"A"£PI"Q21"2 I S23 [Wll] •QOUT I T23 (FA) [Wll] (T20"SG:SC$31)*A"£PI,T21"2 I P 2 A [ W17] 'QCOOL [ " " ] I Q 2 (FA) [Wll] +SG:$CS50"(Q22-Q23) I S2A [Wll] •QCOOL I T2A (FA) [Wll] +SG:SCSA9"(T22-T23) I P25 [ W17 J 'MRPRE [kg" 1/3] I Q25 (FA) [Wll] ($G:SC$AA'Q2A/SG:$C$52) I S25 [Wll] 'MRPRE I T25 (FA) [Wll] (SG:SCSA3,T2A/SG:SCS51) I 027 [W5] ' ROCKET DESIGN DETAILS (stage dependence) I P28 [ W 1 7] 'FEED SYSTEM USED 1 Q28 [Wll] "press. I R28 [Wll] "press. I S 28 [Wll] "press. 1 T28 [Wll] "press. I P29 [ W 17] •THRDES [kg] I Q29 (FO) [Wll] +SG:$C$66"Q86 I R29 f FO) [Wll] +SG:$CS66'R86 I S 29 (FO) [Wll] +SG:SC$66*S86 I T29 (FO) [Wll] +SG:SCS66*T86 I P 30 (F2) [W17] 'TIHEB [sec] I Q30 (F2) [Wll] +Q79"Q13/Q29 I R30 (F2) [Wll] piF(R29«0,0,R79*R13/R29) I S30 (F2) [Wll] eiF(S29«0,0,S79«S13/S29) I T30 (F2) [Wll] gIF(T29«0,0,T79"T13/T29) I P31 [ W17] •OPT. FEED SYSTEM I Q31 [Wll] PIF(Q30<=(8130.l«Q29"-0.75867)," press."," pimp") I R31 [Wll] eiF(R29 =0, " n/a",|IIF(R30<"(B130.1" R29"-0 press. pump" )) I S31 [Wll] gIF(S29»0," n/a",gIF(S30<«(8130.l*S29--0 press. pump" )) I T31 [Wll] gIF(T29»0," n/a",eiF(T30< =(8130.l*T29--0 press. pump" )) I:P32: (W17) 'PT [MPa] I:Q32: (F4) [Wll] §IF(Q28-"PUMP",50«0.101325/14.696,1.5•SC:SFS17"0.101325/14.696 ) I:R32: (F4) [Wll] £IF(R28»"PUMP",50"0.101325/14.696,1.5'SC:SF$17«0.101325/14.696 ) I:S32: (F4) [Wll] $IF(S28»"PUMP",50*0.101325/14.696,1.5*SC:SFS17*0.101325/14.696 ) I:T32: (F4) [Wll] 0IF(T28«"PUMP",50«0.101325/14.696,1.5«SC:$F$17"0.101325/14.696 ) I:P33: [W17] 'PINDEX I:Q33: (F0) [Wll] pIF(Q32>«SD:SI$71,7,gIF(Q32>«SD:SHS71,6,0IF(Q32>=SD:SGS71,5,pi F(Q32>=SD:$FS7L,4,piF(Q32>»$D:$E$7113,SD:SJS61))))) I:R33: (FO) (Wll) 0IF(R32>»SD:SIS71,7,0IF(R32>»SD:SHS71,6,£IF(R32>»SD:SGS71,5,pi F(R32>=SD:SFS71,4,piF(R32>»SD:SES71,3,SD:SJS61))))) I:S33: (FO) [WU] gIF(S32>»SD:SIS71,7,#IF(S32>«SD:SHS71,6,§IF(S32>«SD:SGS71,5,pi F(S32>=SD:SFS71,4,piF(S32>-SD:$E$71,3,$D:SJS61))))) I :T33: (FO) [Wll] piF(T32>«SD:SIS71 ,7,{1IF(T32> = SD:SHS71 ,6,0IF(T32>»SD:SGS71,5,0I F(T32> = SD:SFS7l,4,(iIF(T32> = SD:SES71,31SD:SJS61))))) P34 (W17] 'VOX [n3/kg] Q34 [Wll] pINDEX(SD;SBS71..SD:SIS87,Q33,SD:$FSU5) R34 [Wll] piNDEX(SD:S BS71 ..SD:SIS87,R33,SD:SFS115) S34 [Wll] piS'DEX(SD:SBS71..$D:SIS87,S33.SD:SFS115) T34 (Wll) piNDEX(SD:SB$71..SD:SIS87,T33.SD:SFS115) P35 [W17] •VF [n>3/kg] Q35 [Wll] piS'DEX(SD:SBS71.. SD: S IS87 , Q33 ,SD: SFS 114) R35 [Wll) pi NDEX(SD:SBS71, SD:SIS87,R33,SD:SFS114) S35 [Wll] piN'DEX(SD:SBS71..SD:SIS87,S33,SD:SFS114) T35 [Wll) piNDEX(SD:SB$71..SD:$IS87,T33,SD:SFS114) P36 [wi7] 'oxrc Q36 (F4) [Wll] piF(Q28 ="press.",3"SG:SCS62*SG:SCS84*Q32* 1000000*Q34«(1-Q14)/( 2*SG:SCS85,(1+(1/SC:SFS16))),3"SG:SCS83,SG:SCS84,Q32*1000000*Q34"(1-Q14)/(2"SG:S CS85"(1+(1/SC:SFS16)))) 1 : R36: (F4) [Wll] p IF(R28 • "press ." , 3'SG : SCS82 • S G : SCS84• R32 • 1000000 • R34 "(1 -R14)/ ( 2*SG:SCS85,(1+(1/SC:SFS16))),3•SG:SCS83«SG:SCS84•R32•1000000•R34•(1-R14)/(2•SG:S CS85'(1+(1/SC:SFS16)))) 1 :S 36 : (F4) [ W11J piF(S28«"press.",3 *SG:SCS82*SG:SCS84«S32•1000000'S34*(1-S14)/( 2*SG:SCS85,(1+(1/SC:SFS16))),3"SG:SCS83,SG:SCS84,S32,1000000'S34*(1-S14)/(2*SG:S C$85*(1+(1/SC:$FS16)))) I:T36: (F4) (Wll) piF(T28»"press.",3'SG:SCS82•SC:SCS84«T32•1000000*T34•(1-T14)/( 2,$G:SCS85*(1 +(1/SC:SFS16))),3 *SG:SCS83*SG:SCS84•T32•1000000"T34•(1-T14)/(2•SG:S CS8V(1 +(1/SC:SFS16)))) 1:P 37: [W17) 'FTC I:Q37: (F4) [Wll) piF(Q28®"press.",3"SG:SCS76"SG:SCS78"Q32"1000000"Q35•(1-Q14)/( 2"SG:SCS79"(1+SC:SFS16)),3*SG:SCS77•SG:SCS78"Q32'1000000"Q35•(1-Q14)/(2•SG:SCS79 "( 1 + SC:SFS16))) 1 : R37: (F4) (Wll] piF(R28»"press.",3,SG:SCS76"SG:SCS78*R32"1000000*R35*(l-R14)/( 2*SG:SCS79*(1 + SC:SFS16)),3 *SG:SCS77•SG:SCS78'R32•1000000*R35*(1-R14)/(2*SG:SCS79 *(1 + SC:SFS16 ))) I :S37: (F4) (Wll] piF(S28«"press.",3"SG:SCS76,SG:$CS78"S32,1000000"S35"(l-S14)/( 2*SG:SCS79"(1 + SC:SFS16)),3"SG:SCS77"SG:SCS78*S32"1000000"S35•(1-S14)/(2 *SG:SCS79 "(1+SC:SFS16))) 1:T37: (F4) (Wll) piF(T28«"press.",3•SG:SCS76-SG:SCS78«T32•1000000*T35•(1-T14)/( 2"SG:SCS79"(1+SC:SFS16)),3"SG:SCS77«SG:SC$78*T32"1000000"T35*(1-T14)/(2"SG:SCS79 *(1 + SC:SFS16))) I:P38: [W17] ' FSC I:Q38: [Wll] §IF(Q28="press.",SG:SCS29*SC:SFS17/300,SG:$C$30*pEXP(0.000759*SC:$F S17)) I:R3B: (Wll) $IF(R28="press.",$G:$CS29*SC:SFS17/500,SG:$C$30*gEXP(0.000759*SC:SF S 17 )) 1 :S 38: [Wll] PIF(S28«"press.",SG:SCS29*SC:SFS17/500,$G:$CS30*gEXP(0.000759*$C:$F S17)) I:T30: [Wll] PIF(T28«"press.",$G:SC$29*$C:SF$17/500,SG:SC$30*gEXP(0.000759*SC:$F 517)) OAO [W5J • INTERMEDIATE MASS CALCS PA1 [W17] •COEFSUM QA1 (FA) (Wll) +SG:$C$35+SG:SCSA8+Q37+Q36+SG:$CS59+$G:SCS60+Q38+SG:SCS22 RA1 (FA) (Wll) +SG:SCS35+SG:$CSAB+R37+R36+$G:SCS59+SG:SCS60+R38+SG:SCS22 SA1 (FA) Ikfllj +SG:SCS35 +$G:SCSA8+S37+S36+SG:SCS59+SG: SCS60+S38+SG:$CS22 TA1 (FA) (Wll) +SG:SCS35+SG:SC$A8+T37+T36+SG:SCS59+SG:SCS60+T38+SG:SCS22 PA2 I W17] 'COEFTL QA2 (FA) (Wll] @IF(T7»"YES"WAND"T6»1,SG:SCS1A+QA1,QA1) RA2 (FA) (Wll] 0IF(T7«"VES"«AND#T6»2,SG:SCS1A+RA1,RA1) S i. 2 (FA) [Wllj f)IF(T7 ="YES"«AND«T6»3,SGISCS1A+SA1.SA1) TA 2 (FA) (Wll] PIF(T7="YES"«AND«T6*A,SG:SCS1A+TA1,TA1) P^i 3 [W17] 'COEFTLM QA3 (FA) (Wll) PIF(Q7«"N0" ,0,(11 F(T7 = " YES"#AND«T6*1.SGi SCS1A+SG:SCS35+SG:SCS60_ , SG; SCS35+ SG : SCS60)) I:RA 3: (FA) (Wll] gIF(Q7»"N0",0,eiF(T7»"YES"«AND«T6»2,SG:SCS1A+SG:SCS35+SG:SCS60 ,SG:SCS 35 + SG : SCS6D)) I :SA3: (FA) (Wll] piF(Q7="N0",0,^IF(T7="YES"«AND«T6=3,SG:SCS1A+SG:SCS35+SG:SCS60 , SG:SCS35+SG :SCS60)) I :TA3: (FA) (Wll] gIF(Q7="NO",0,eiF(T7»"YES"«ANDffT6=A,SG: SCS1A+SG:SCS35+SG:SCS60 I P A A [W17] 'NUMER I QA A (FA) (Wll) +Q1A-QA2 I RA A (FA) (Wllj +R1A-RA2 I S A A (FA) (Wllj +S1A-SA2 1 T A A (FA) (Wllj +T1A-TA2 1 PA5 (W17 1 'DEN OK 1 QA5 (FA) (Wll] +SG:SCS56"(1 -Q1A) I R*5 (FA) (Wll) +SG:SCS56*(1-R1A ) I SA 5 (FA) (Wll) +SG:SCS56-(1-S1A) ] T A 5 (FA) (Wll) +SG:SCS56*(1-T1A ) 1 PA6 (W17 ] 'VALUE I QA6 (FA) [Wll] 100*QA A/QA5 I RA 6 (FA) (Wll] §IF(RA5=0,0,100*RAA/RA5) I S A 6 (FA) (Wll) 0IF(SA5»O,O, 100-SAA/SA5) 1 TA 6 (FA) [Wll] 0IF(TA5-O,O,100*TAA/TA5) I PA 7 (W17] 1 VAL2 I QA7 (FA) (Wll) 1-QA 2 1 RA 7 (FA) (Wll) 1-RA2 I SA7 (FA) (Wll] 1-SA 2 1 TA 7 (FA) (Wll) 1-TA2 I PAS (W17) 1 V2HL 1 QA8 (F2) (Wll] +QA7*Q11 I RA8 (F2) (Wll] +RA7 * R11 I SAB (F2) (Wll) +SA7*S11 1 TAB (F2 ) (Wll) +TA7*T11 1:050: [W5) ' SOLVING CUBIC FOR MF I:P51: [W17] 'A1 I:Q51: (S3) [Wll] 0IF(Q8«"No"#AND#T8«"No",-(3»Q I T73 (F2) [Will + SG:SC$ii7"(T86-T61i) I P74 [W17] 'MFT [kg] I Q7 U (F2) [Wll] +Q37"(Q86-Q6'i) I F74 t F2) [Wll] + R37*(R86-R6'i) I S74 (F2) [Wll] +S37* (S86-S6 I AK16 [W17] 1 SFB [W/n>2] I AL16 {Wll] +SG:SCS61'(SG:$CS15/AL15)"2 I AM16 [Wll] +SG:SCS61"(SG:SCS15/AH15)-2 I AN 16 (Wll] +SG:SC$61*($G:$CS13/AN15)*2 I A016 [Wll] +SG:SCS61"(SG:SCS15/A015)*2 I AJ10 [W5] ' REFRIGERATION MASS PRELIMINARY CALCULATIONS I AK19 [W17 ] ' Oxidizer I AN 19 [Wll] ' Fuel I AK20 (W17] 'HOX [W/n2] 1 AL20 (FA) [Wll] +SG:SCS58"$G:SCS21*SG:$CS6A-A I AS'20 [Wll] 'HF I A020 (FA) [Wll] +$G:SC$58»SG:$C$20'SG:$CS62*A 1 AK21 [W17 ] 'ROXT [m/(kg)"1/3] I AL21 (F"£<) [Wll] (0.75/(SG:SCS18"SG:SCS50'gPI))-(l/3) 1 A N 21 [Wll] ' RFT I A021 (FA) [Wll] (0.75/(SG:SCS16"SG:SCSA9"pPI))*(l/3) I AK22 [*>17] 'QIN (W/(kg)"2/3] I AL22 (F*.) (Wll] (AL16*(1-SG:SCS5A)«SG:SCS2A)•gPI'AL21 "2 I AN 2 2 [Wll] 'QIN I A022 (FA) [Wll] (AO16"(1"SG:SCS3A)*SG:SCS23)*0PI"AO21 "2 I AK 2 3 [W1 7 ] ' QOl'T [ " " ] I AL23 (FA) [Wll] (AL20"SG:SCS32),A"@PI,AL21*2 I AN 2 3 [Wll] ' QOI T I A023 (FA) [Wll] (A020'SG:SCS31)"A"pPI"A021"2 I AK2A [ K1 7] 'QCOOL [ " " ] I AL2 A (FA) (Wll] +SG:SCS50"(AL22-AL23) I AN2A [Wll] 'QCOOL I Atm (FA) (Wll] +SG:$CSA9*(A022-A023) I AK25 [ W17] 'KRPRE [kg" 1/3] I AL2S (FA) (Wll] (SG:SC$AA"AL2A/SG:SCS52) I AN 25 [Wll] 'MRPRE I A025 (FA) (Wll] (SG:SCSA3"A02A/SG:SCS51) I AJ27 [V.5 ] ' ROCKET DESIGN DETAILS (stage dependence) 1 AK 28 [ W 17 ] 'FEED SYSTEM USED I AL28 [Wll] "Dunp I A" 2 8 [Wll] "pump I AN 28 [Wll] "pump I A02B (Wll) "pump 1 AK29 [ W 1 7] 'THRDES [kg] I AL 29 (F0) (Wll] +SG:SCS66"AL86 I AC.29 (FO) [Wll] +SG:SCS66"AM86 I AN 29 (FO) (Wll) +SG:SCS66*AN86 I A029 (FO) [Wll] +SG:SC$66*A086 I AK30 (F2) [W17] 'TIMEB [sec) I AL30 (F2) [Wll] +AL79*AL13/AL29 I AM30 (F2) [Wll] gIF(AM29"0,0,AM79"AM13/AM29) I AN 30 (F2) [Wll] gIF(AN29»0,0,AN79*AN13/AN29) I A030 (F2) [Wll] gIF(A029«0,0,A079,A013/A029) I AK31 [W17 ] 'OPT. FEED SYSTEM I AL31 [Wll] PIF(AL30<»(8130.1"AL29"-0.75867) ," press." , pump") I AC.31 [Wll] £IF(AM29=0," n/a",0IF(AM30<»(8130.1" AM29 pres s " I " pump'')) 1 AN31 [Wll] gIF(AN29=0," n/a",gIF(AN30<«(8130.1* AN29 pres s • puop")) 1 A031 [Wll] gIF(A029=0," n/fl",gIF(A030<»(8130.1" A029 pres s M 1 •• purap")) I:AK32: (W17] 'PT (MPa] I:AL32: (FA) (Wll) $IF(AL28« •PUMP",50*0.101325/1 A.696,1.5*SC:SFS8*0.101323/1A.69 6) I:AM32: (FA) [Wll] £IF(AM28»"PUMP",50*0.101325/1A.696,1.5*$C:$FS8*0.101325/14.69 6) I:AN 32: (FA) [Wll] gIF(AN28= "PUMP"|50*0.101325/1A.696.1.5*$C:$FS8"0.101325/1A.69 6) I: A032: (FA) [Wll] £IF(A028 "PUMP",50*0.101325/1 A.696,1.5»$C:$F$8«0.101325/1A.69 6) I:AK33 .• [W17] PINDEX I: AL33: (F0) [Wll] @IF(AL32>,,$D:$1$71,7,@IF(AL32>°$D:$H$71,6,@IF(AL32>B$D:$G$71, 5,eiF(AL32>«$D:SF$71,A,0IF(AL32>«$D:$E$71,3,SD:SJS61))))) I.AH33: (FO) [Wll] §IF(AM32>=SD:SIS71,7 .glF(AM32 >«SD:SHS71, 6 ,£IF(AM32> • SD: SGS71, 5.^IF(A.r32> = SD: SFS71 tA ,^IF(AM32>-SD:SES71,3,SD: SJS61))))) I:A X 33 : (FO) [Wll] glF(AN 32>«SD:S1$71,7,0IF(AN32>»$D:SHS71,6.gIF(AN 32> = SD:SGS71, 5,gIF(AN32>=SD:SFS71,A,gIF(AN32>«SD:SES71,3,SD:SJS61))))) 1:A033: (FO) [Wll] gIF(A032>=$D:SIS71,7,gIF(A032>-SD:SHS71,6,gIF(A032>=SD:SGS71, 5,gIF(A032>«SDiSFS71,A,gIF(A032>-SD:SES71,3,SD:SJS61))))) 1 :AK3A [W17] •VOX [m3/kg] ]: AL3A [Wll] gINDEX(SD:SBS71..SD:SISB7,AL33,SD:SFS113) I :A"3A [Will gINDEX(SD:SBS71..SD:SIS87,AM33,SD:SFS113) AN 3 A [WU] g I S'DEX( SD :S BS 71. SD:SIS87,AN33,SD:SFS113) A03A [Wll] gINDEX(SD:SBS71. SD:S1S87,A033,SD:$F$113) A K 3 5 [W17] • VF [ra3/kg] AL35 [Wll] gINDEX(SD:SBS71..SD:SIS67,AL33,SD:SFSU2) AV.35 [Wll] gISDEX(SD:SBS71. ,SD:SIS87,AM33,SD:SFS112) A!.'35 [Wll] gl NDEX(SD:SBS71. .SD:SIS87,AN33,SD:SFS112) I : A035 [Wll] g t S'DEX(SD:SBS71. , SD:SIS87,A033,SD:SFS112) I : AK36 [W17] ' OXTC 1 : A L 36 (FA) [Wll] piF(AL28="press ",3'SG:SCS82"SG:SCS8A-AL32*1000000'AL3A"(1-AL 1A)/(2,SG:SCS05,(1+(1/SC:SFS7))),3«SG:SCS83'SG:SCS8A*AL32*1OOOOOO*A13A*(1-AL1A)/ (2*SG:SCS85*(1+(1/SC:SFS7)))) I : A u. 36 : (FA) [Wll] gIF( AK28 = "press 3 * SG :SCS62 • SG: SCS8A *AM32 * 1000000• AM3A *( 1-AM 1A)/(2"SG:SCS85*(1 +(1/SC:SFS7))),3 *SG:SCS83•SG:SCS8A•AM32* 1000000'AM3A•(1-AM 1A)/ (2"SG:SCS85"(1+(1/SC:SFS7)))) I : A N 36 : (FA) [Wll] (31F(AN28 = " pres s." , 3• SG:SCS 82 • SG : SCS8A * AN32 * 1000000 * A N3A *(1 - AN 1A)/(2*SG:SCS05"(1+(1/SC:SFS7))),3"SG:SCS83*SG:SCS8A*AN32*1OOOOOO*AN3A"(1-AN1A)/ (2'SG:SCS85"(1+(1/SC:SFS7)))) I:A03fi: (FA) [Wll] gIF(A028="press3*SG:SCS82•SG:SCS8A*A032* 1000000"A03A"(1-AO 1A)/(2"SG:SCS85,(1+(1/SC:SFS7))),3*SG:SCS83*SG:SCS8A*A032*1000000*AO3A*(1-A01A)/ (2'SG:SCS85»(1+(1/SC:SFS7)))) I:AK37: [W17] 'FTC I : AL37: (FA) [Wll] gIF(AL28="press3*SG:SCS76*SG:SCS78*AL32* 1000000*AL35*( 1 -AL 1A)/(2"SG:SCS79"(1+SC!SFS7)),3"SG:SCS77*SG:SCS78*AL32*1000000*AL35*(1-AL1A)/(2*S G:SCS79*(1+SC:SFS7))) I : A M 37: (FA) [Wll] gIF(AM28«"press3*SG:SCS76*SG:SCS76-AM32* 1000000"AM35*(1-AM 1A)/(2*SG:SC$79*(1+$C:SFS7)),3«SG:SCS77"SG:SCS76*AM32"1000000*AM35*(1-AM1A)/(2*S G:SCS79*(1+SC:SFS7))) I : AN37: (FA) [Wll] gIF(AN28="press.",3*SG:SCS76'SG:SCS78*AN32*1000000*AN35*(1-AN 1A)/(2*SG:SCS79*(USC:SFS7)),3*SG:SCS77*SG:SCS78*AN32*1000000*AN35*(1-AN1A)/(2*S G:SCS79*(1+SC:SFS7))) I : AO 37: (FA) [Wll] gIF(A028= "press.",3*SG:SCS76*SG:SCS78"A032* 1000000*A035*(1-AO 1A)/(2*SG:SCS79*(1+SC:SFS7)),3*SG:SCS77*SG:SCS78*A032*1000000"A035*(1-A01A)/(2»S G:SCS79*(1+SC:SFS7))) I:AK38 IW17] 'FSC I: AL38 (FA) (Wll) #IF(AL28*"press. ',SGJSCS29"SC:SFS8/500,SG:SCS3O»0EXP(0.000759 "SC:SFS8)) I:AM38 : (FA) [Wll] #IF(AM28«"press. '.SG:SCS29*$C:SFS8/500,$G:SCS30»£EXP(0.000759 •SC:SFS8)) I: AN38: (F^t) [Wll] §lF(AN28«"press. ',$G:$C$29*$C:$F$8/500t$G:SCS30"{IEXP(0.000759 *SC:SFS8)) I:A038: (FA) [Wll] 0IF(A028«"press.SG:SCS29*$C:SFS0/5OO,SG:SCS30"0EXP(0.000759 •SC:SFS8)) : AJAO [W5] * INTEKMEDIATE MASS CALCS :AK*. 1 [W17] 'COEFSUM :ALA1 (FA) (Wll] +$G:SCS35+SG:SC$A7+AL37+AL36+SG:SCS59+SG:SC$60+AL38+SG:SC$22 :AMA1 (FA) [Wll] +SG:SC$35+SG:SC$A7+AM37+AM36+SG:SCS59+SG:SCS60+AM38+$G:SCS22 :ANAI (FA) [Wll] +SG:SCS35+SG:$CSA7+AN37+AN36+SG:SCS59+SG:SCS60+AN38+SG:SCS22 :A0A1 (FA) (Wll] +$G:SCS35+SG:SCSA7+A037+A036+SG:SCS59+SG:SCS60+A038+SG:SCS22 :AKA2 IW17] 'COEFTL :ALA 2 (FA) (Wll) £IF(A07= YES"»AND»A06«1,SG:SCSIA + ALA1,ALA 1) :AMA2 (FA) [Wll] £IF(A07= YES"«AND«A06=2,SG:SCSIA + AHA 1,AHA 1) : ANA 2 (FA) [Wll] gIF(A07= YES"#ANDWA06»3,SG:SCSIA + ANA 1,AM A1) I: AO^ 2 (FA) [Wll] $IF(A07» YES"WANDl/A06«A ,SG: SCS1A+A0A1 , A0A1) I :AKA 3 (W17] 'COEFTLM I : AL'i 3 (FA) (Wll] gIF(AL7= NO' ,0,@IF(A07="YES"«AND»A06=1,SG:SCS1A + SG:SCS35+ SG:S, CS60.SG :SCS35+SG:SCS60)) I : AMA3: (FA) (Wll] gIF(AL7= 'NO' ,0,gIF(A07=" YES "//AND«A06»2, SG:SCSI A + SG :SCS35 + SG : S CS60.SG :SCS35 + SG:SCS60)) I IAS')3: (FA) [Wll] glF(AL7= NO" ,0,(UF(A07 = "YES"»AtiD«A06»3, SG:SCS1A + SG:SCS35+ SG : S CS60, SG:SCS35+ SG:SCS60)) I:A0<<3: (FA) (Wll] gIF(AL7= NO" ,0,piF(A07n"YES"tfAND«A06»A,SG:SCS1A + SG:SCS35+SG: S CS60,SG:$C$35+SG:SCS60)) I AKA4 (W17] •NUf.ER I AL'i A (FA) (Wll] +AL1A-ALA2 I A^Ai. (FA) [Wll] +AM1A-AMA2 I ASIA (FA) (Wll) +AN1A-ANA2 . I AO 302 I: AJ50: [W5] ' SOLVING CUBIC FOR MF I:AK51: [W17J 'A1 I.-AL51: (S3) [Wll] gIF(AL6« "No "//ANDWA08* "No",-(3*AL48/AU6) ,QIT (AL8 = "No" ,- ( 3" AU 8/AL'i6+(A025/AL'<6)*3) ,0IF (A08-" No",- (3 • AL48/AU6+ ( ( SC ! $FS7) * ( 2/3) " AL25/AL46)" 3), -(3*AL't8/AL'i6+((A025+(SC:SFS7)'(2/3)*AL25)/AL't6)"3)))) I: AM51 : (S3) [Wll] (> IF( AM46»0 ,0 ,gIF( AL8»" No"0AND#AO8» "No " ,-(3"AM AK71 [W171 •HSE [kg] AL71 I C80 [Will •SCSI I A91 [ W5] 'MARS SURFACE DETAILS I A92 tW5] \ • I B92 IW17) I C92 [Wll] I D92 [Wll] \» I E92 (Wll) \" I F92 (Wll) \» I B93 [ W17) 'MARS ROVER (kg) I C93 [Wll) +SG:SCSA2 I B9A [W17) 'OX. FACTORY (kg) I C9A (F2) (Wll) +SE:SCS31 I D9A [Wll] > I E9A [Wll] •MASS USED I F94 (F2) (Wll) £IF(SXS8»"Yes",C94,0) I B95 [W17 ) 'FUEL FACTORY (kg) I C95 (F2) [Wll] 1008 I D95 (Wll) > I E95 (Wll) 'MASS USED I F95 (F2) [Wll] eiF(SAA$B*"Yes",C95,0) I B97 [W17 ] 'OXIDIZER REFRIGERATION UNITS [kg) I C98 [Wll] "STAGE I I D98 [WU] "STAGE 11 I E98 [Wll] "STAGE 111 I F90 [Wll] "STAGE IV I B99 IW17 ) "LEG 6 I C 99 (FA) [Wll] +SF:SDS116 I D99 (FA) [Wll] +SF:S DS 115 I E99 (FA) [Wll] + SF :SDS114 I F99 (FA) [Wll] +SF:SDS113 I BlOO [W 1 7] -LEG 5 I ClOO (FA) [Wll] +SF:SDS120 I DlOO (FA) [Wll] +SF:SD$119 I ElOO (FA ) [Wll] +SF:SDS118 I FlOO (FA) (Wll) +SFISDS117 I B101 [W 1 7] "LEG A I C101 (FA) [Wll] +SF:SDS124 I D101 (FA ) [Wll] +SFISDS123 I ElOl (FA) [Wll] +SF:SDS122 I FlOl (FA) [Wll] +$F:SDS121 I B103 (W17] "TOTAL I C103 (FA) [Wll] gSUM(C99..FlOl) I D103 [Wll] > > I E103 [Wll] •MASS USED I F103 (FA) (Wll) £IF(SXS7»"Yes" ,C103,0) 1 BIOS [W1 7] 'FUEL REFRIGERATION UNITS (kg) I C106 [Wll] -STAGE I I D106 (Wll) -STAGE II I E106 [Wll] "STAGE III I F106 [Wll] -STAGE IV I B107 [ W17] -LEG 6 I CI07 (FA) (Wll) +SF:SES116 I D107 (FA) [Wll] +SF:SES115 I E107 (FA) [Wll] +SF:SES114 I F107 (FA) [Wll] +SF:SES113 306 Bioe [Ml ] •LEG 5 C108 (F4 [HI 1] +SF:SES120 Dioe (F4 [HI 1] +SF:SES119 E108 (F4 [HI 1] +SF:SES 118 F108 (F4 [H11] +SF:SES117 B109 [W171 -LEG 4 C109 t F4 (*11] +SF:SES124 D109 (F4 (Will +SF:SES123 E109 (F4 [HI 1] +SF:SES122 F109 (F4 [HI 1 ] +SF:SES121 Bill [W171 •TOTAL Clll (F4 [Nil] PSUHCC107,,F109) Dill [W1 ] . > E111 [ w 1 •MASS USED Fill (F4 [HI 1] 0IF(SAAS7= •Yes",Clll,0) B11 3 [ 1 ) •OXIDIZER NECESSARY FOR RETURN C114 [W1 1 "STAGE I D11A t 1 ] "STAGE II E114 [ w ] "STAGE III F114 [W1 ] -STAGE IV BUS {W1 ] -LEG 6 C115 C F4 [Wll] +C78 D115 (F 4 (W11] +D78 El 15 [F4 [Hll] + E78 F11 5 ( F 4 [Wll] +F78 B116 [W1 ] "LEG 5 C116 ( F 4 [Wll] +J78 D116 ( F 4 [Wll] +K78 E116 ( F 4 [Wll] +L78 F116 (F4 [Wll] +M78 B11 7 [ W1 ] "LEG 4 C 117 C F4 [Wll] +Q78 D117 ( F 4 [Wll] +R78 El 17 (F4 [Wll] +S78 F 117 ( F 4 [Wll] +T78 B1 1 9 E W17) "TOTAL C 1 1 9 ( F 4 [Wll] £SUM(CI15. .F117) D119 [W1 ] > E119 [W1 ] •MASS USED F119 (F4 [Wll] @IF(SXS8«" Yes".CI 19,0) B1 2 1 [HI ] •FUEL NECESSARY FOR RETURN [kg) CI 22 [ W1 1 -STAGE I D1 2 2 [ w J "STAGE II E122 [HI 1 -STAGE III F122 [HI ] -STAGE IV B123 (H171 -LEG 6 C123 (F4 [Wll] +C77 D123 ( F 4 [Hll] +D77 El 23 (F 4 [Hll] +E77 F1 23 (F4 [Wll] +F77 B124 (HI ) "LEG 5 Cl» (F4 [Hll] +J77 D124 (F4 [Hll] +K77 El 24 (F4 [Hll] +L77 F124 (F4 [Hll] + M77 307 I:B125: [W17] "LEG A I:CI25: (FA) [Wll] +Q77 I:D125: (FA) [Wll] +R77 I:El25: (FA) [Wll] +S77 I:Ft 25: (FA) [Wll] +T77 I:B127: [W17] "TOTAL I:C 127: (FA) [Wll] #SUK(C123..F123) I:D127: [Wll] ' > I:E127: [Wll] 'MASS USED I : F1 27: (FA) [Wll] piF($AA$8«"Yes",C127,0) H89: [W5) 'TOTAL MASS SUMMARY J89: [Wll] +SCS1 H91: [W5] ' PRE-MARS K91: (Wll] ' POST-MARS 192: [W17] 'MA (kg] J92: (FO) [Wll] @SUM(X63..AA63)+gSUM(AE63. .AH63)+@SUM(AL63. . A063) L92: [Wll] 'MA [kg] M92: (Fl) (Wll] $SUM(C63..F63)+gSUM(J63..M63)+$SUM(Q63..T63) 193: (W17] 'MG [kg] J93: (FO) [Wll] £SUM(X65..AA65)+£SUM(AE65..AH63)+$SUM(AL65..A065) L93: [Wll] 'MG [kg] M93: (Fl) [Wll] §SUM(C65..F65)+gSUM(J65..M65)+gSUM(Q65..T6J) 194: {W171 'MSS (kg] J94: (FO) [Wll] £SUM(X66..AA66)+£SUM(AE66..AH66)+£SUM(AL66..A066) L94: [Wll] 1MSS [kg] M94: (Fl) (Wll] 0SUM(C66..F66)+§SUM(J66..M66)+gSUM(Q66..T66) 195: [W17] 'HRF [kg] J95: (FO) [Wll] flSUM(X67..AA67)+gSUM(AE67..AH67)+^SUM(AL67..A067) L95: [Wll] 'MRF [kg] "95: (Fl) [Wll] 0SUM(C67..F67)+0SUM(J67..M67)+0SUM(Q67..T67) 196: [W17] ' MROX [kg] J96 : (FO) [Wll] gSUM(X60..AA60)+0SUM(AE68..AH68)+PSUM(AL68..A068) L96: [Wll] 'MROX [kg] M96: (Fl) [Wll] gSUM(C68..F68)+gSUM(J68..M68)+£SUM(Q68..T68) 196: [W17] 'ME [kg] J98: (FO) [Wll] @SUM(X70.. AA70)+@SUM(AE70. .AH70)+gSUH(AL70. .A070) L98: [Wll] 'ME [kg] M98: (Fl) [Wll] £SUM(C70. . F70) +PSUM( J70 . . M70)t-gSUM (Q70. . T70) 199: [W1 7] 'MSE [kg] J99 : (FO) [Wll] ()SUM(X71. . AA71)+0SUM(AE71 .. AH71)+£SUH(AL71. .A071) L99: [Wll] 'MSE [kg] M99 : (Fl) (Wll] (ISUM(C71.. F71) +0SUM (J71. . M71)+£SUM (Q71. . T71) IlOO: [W17] 'KFS [kg] JlOO: (FO) [Wll] @SUM(X72..AA7 2)+0SUM(AE72..AH72)+0SUM(AL72..A072) L100: [Wll] 'MFS [kg] '.100: (Fl) [Wll] 0SUM(C72..F72)+0SUM(J72. .M72)+#SUM(Q72. .T72) 1101: (Vf 17] 'MN [kg] J 101: (FO) [Wll] £SUM(X73..AA73)+PSUM(AE7 3.. AH73)+£SUH(AL73..A073) L101: [Wll] 'MN [kg] «101 : (Fl) [Wll] £SUM(C73. .F73)+fiSUM( J73..M73)+flSUM(Q73. .T73) 1102: [W17] 'MFT [kg] J102 : (FO) (Wll] §SUM(X74..AA74)+gSUM(AE74.,AH7A)+@SUM(AL74..A074) L102: [Wll] 'MFT [kg] 1102: (Fl) (Wll) @SUM(C74..F7'i)+gSUM(J74. .M74)+£SUM(Q7 R88 [Wll] +SCS1 090 [N5) ' TOTAL MISSION P91 [W17 ] 'MA [kg] Q91 (F2) (Will +J92+M92 591 [Will 'ME [kg] T91 (F2) [111] +J98+M90 P92 [W17] 'MG [kg] Q92 (F2) [Will +J93+H93 592 [Wll] 'MSE [kg] T92 (F2) [W11J +J99+H99 P93 [W17) 'MSS [kg] Q93 (F2) [Wll] +J94+M94 593 [Wll] 1MFS [kg] T93 (F2) [Will +J100+M100 P9<< [W17] 'MRF [kg] Q9 B16 [ W17 'SFB [W/»2] C16 [Wll +$G:$D$6l*(SG:$DS15/C15)"2 D16 [Wll +SG:SD$61'(SG:SDS15/D15)"2 E16 [Wll +SG:SDS61"($G:SDS15/E15)"2 F16 [Wll +$G:$D$61"(SG:SD$15/F15)*2 A1G [W5J • REFRIGERATION MASS PRELIMINARY CALCULATIONS B19 [W17 ' Oxidizer EL9 (Wll • Fuel B20 [W17 'HOX [W/b2] C20 (FA) (WUJ +SG:SDS58*SG:$DS21*SG:SDS65"A E20 [Wll •HF F20 (Ft) [Wll] +$G:$D$58*SG:$D$20*$G:$D$63"«SD:SIS71,7,gIF(C32>"SD:$H$71t6,gIF(C32>»SD:SG$71,5lgI F(C32>=SD:SFS71,4,gIF(C32>»SD:SES71,3,SD:SJS61))))) J:D33: (FO) {Wll] £IF(D32>°SD:$I$71,7,£IF(D32>'SD:$H$71,6i£IF(D32>°$D:$GS71,5,£I F(D32>=SD:$FS71,4,gIF(D32>»SD:SES71,3,$D:$JS61))))) J:E33: (FO) [Wll] piF(E32>*SD:SIS71,7,gIF(E32>»SD:SHS71,6,gIF(E32>«SD:SGS71,5,§I F(E32>*SD:SFS71,4,#1F(E32>»SD:$ES71,3,$D:$JS61))))) J:F33: (FO) [Wll] glF(F32>=SD:S1$71,7,gIF(F32>=SD:SHS71,6,gIF(F32>=SD:$G$71,5,gl F(F32>=$D:$FS71,4,gIF(F32>«$D:SE$71,3,SD:SJS61))))) :B34 {W17) 'VOX [«3/kg] •-C3U [Wll] piNDEX(SD:SBS71..SD:SI$07,C33,SD:SFS120) :D34 [Will gINDEXtSD:SBS71..SD:SIS87,D33,SD:SFS120) :E3<. (Wll] gINDEX(SD:SBS71..SD:SIS87,E33,SD:$FS120) : F34 [Wll) gINDEX(SD:SBS71..SD:SI$87,F33,SD:SF$120) :B35 IK17] 'VF [o3/kg] : C35 [Wll) piNDEX(SD:SBS71..SD:SIS87,C33,SD:SFS119) :D35 [Wll) glNDEX(SD:SBS71..SD:SIS87,D33,SD:SFS119) : E35 [will gINDEX(SD;SBS71..$D:SIS87.E33,SD:SFS119) :F35 [Wll] gINDEX(SD:SBS71..SD:SIS87,F33,SD:SFSU9) J: B36 [W17] 'OXTC J:C36: (F4) [Wll] glF(C28""press.gIF(C28-"press.",3'SG:SDS82"$G:SDS84*C32•1000000*C34*(1-C14)/( 2,SG:SDS85"(1 +(1/SC:SFS36))),3*SG;SD$83"SG:SDS84»C32*1000000*C34*(1-C1'.)/(2*SG:S DS85*(1+(1/SC:SFS36)))) J:D36: (F4) [Wll] gIF(D28= "press.",3"SG:SDS82*SG:SDS84*D32* 1000000"D34•(1-D14)/( 2*SG:SDsOS'(l +(i/SC:SFS36))),3*SG:SDS83*SG:SDS84*D32*1000000*D34•(1-D14)/(2*SG:S DS85Ml +(1/SCISFS36)) ) ) J : E36: (FM [Wll] gIF (E28 « "pr es s . " , 3 * SG :SDS82 * SG: S DS84*E3 2 * 1000000*E34• (1 -El 4) / ( 2»SG:SDS85-Cl+<1/SC:SFS36))),3*SG:SDS83*SG:SDS84*E32*1000000*E34*(1-E14 )/(2*SG:S DS85"(1+(1/SC:SFS36)))) J:F36: (F4) [Wll] glF(F28•"press.",3*SG:SDS82*SG:SD$84*F32* 1000000*F34•(1-F14)/( 2*SG:$DS85"(1+(1/SC:SF$36))),3*SG:SD$83*$G:SDS84*F32*1000000*F34*(1-F14)/(2*SG:$ DS85"(1+(1/SC:SFS36)))) J:B37: J W 171 'FTC J:C37: (F4) [Wll] gIF(C28="press.",3"SG:$D$76*SG:$DS78*C32*1000000*C35*(l-C14)/( 2*SG:SDS79*(1+SC:SFS36)),3*SG:SDS77•SG:SDS78*C32*1000000*C35*(1-C14)/(2 *SG:SDS79 •(1+SC :SFS36))) J:D37: (F4) [Wll] gIF(D28»"press.",3*SG:SDJ76*SG:SD$78*D32*1000000*D35*(1-D14)/( 2*SG:SDS79*(1+SC:$FS36)),3*SG:SDS77*SG:$D$78*D32*1000000*035*(1-D14)/(2*$G:SDS79 "(1+SC:SFS36))) J:E37: (F4) [Wll] gIF J B60 |W17J •MF [Kg] J C60 (F J E7 3 F2) [till] +$G:SD$47*(E86-E64) J F73 F2) [Wll] +$G:SDS47"(F86-F64) J B7 [Wll] + E77'SC:SFS 36 J F?0 F2) [Wll] +F77"SC:SFS36 J B79 *17] 'HP [kg] J C79 F 2) [Wll] +C77+ C70 J D79 F2) [Wll) +D77+D78 J E79 F2) [Wll] +E77+E78 J F79 F2) [Wll] +F77+F78 J B81 W17] 'NUHMOD J B82 W 17 ] 'MMODTL [kg] J B83 W17] 'MODTHRTL [kg] J B85 kl7] 'y.l [kg] J C85 F 2) [Wll] @SL'M(C63.. C 75 ) J D85 F 2 ) [Wll] £SUM(D63. .D75) J E85 F 2 ) [Wll] gSl'H( E63. . E75 ) J F85 F 2 ) [Wll] $SUK(F63..F75) J Be& 1.17) 'MO [kg] J C86 F2) [Wll] +C79/(1-C1A ) J D86 F 2) [Wll) §]F(D1W,0,D79/(1-D14)) J E86 F2) [Wll] §IF(E1W.0,E79/(1-El J 116 [W17] •SFB [W/m2] J J1 6 [Wll] +SG:SDS61*(SG:SD$15/J15)-2 J K16 (Wll) +$G:$DS61*(SG:$DS15/K15)"2 J LI6 [Wll] +$G:SD$61*(SG:SDS15/L15)"2 J M16 [Wll] +$G:SDS61*($G:$D$15/M13)-2 J H18 (W5) • REFRIGERATION MASS PRELIMINARY CALCULATIONS J 119 [W17] ' Oxidizer J L19 (Wll] • Fuel J 120 [W17] 'HOX [W/m2] J J20 (F1) (Wll) +$G:$D$38*$G:$D$21*$G:$D$6S*4 J L20 [Wll] •HF J M20 (F«) (Wll) +$G: $DS58*$G:$D$20*$G:$D$63-"$D:SIS71,7,gIF(J32>»SD:SHS71,6,piF(J32>«SD:SGS71,5,gI F(J32>»SD:$FS71,4,piF(J32>-$D:$ES71,3,$D:SJS61))))) J:K33: (FO) (Wll) PIF(K32>«SD:SIS71,7,0IF(K32>«SD:SHS71,6.gIF(K32>=SD:SGS71,3,gl F(K32>-SD:SF$71,4,piF(K32>«SD:SES71,3,$D:SJS61))))) J:L33: (FO) (Wll) @IF(L32>*SD:SI$71,7,@IF(L32>*$D:$H$71,6,0IF(L32>>SD:$G$71,5,@I F(L3 2>=SD:$FS71,4,£IF(132>«SD:SES71,3,SD:$J$61))))) J: M33: (FO) (Wll) 0IF(M32> = S D:SIS71,7,0IF(M32>-$D:SHS71,6,^1F(M32>«SD:SGS71,5,$I F(M32>=SD:SFS71,4,0IF(M32>*SD:SES71,3,SD:SJS61))))) 134 IW17) 'VOX [a3/kg] J34 (Wll) piNDEX($D:$B$71 •SD:S1S87,J33,$D:$F$120) K34 (kill] 0INDEX(SD:SBS71 ,SD:SI$87,K33,SD:SF$120) L3T (Wll) 0ISDEX(SD:SBS71 ,SD:SIS87,L33,SD:$FS120) M 3^ (Wll) 0INDEX(SD:SBS71 •SD:SIS87,M33,SD:SFS120) 135 I W17 J 'VF [n3/kg] J35 (Will eiNDEX(SD:SBS71 .SD:SIS87,J33,SD:SFS 119) K35 (Will @1NDEX(SD:SBS71 .SD: SIS 87,K33.SD:$FS119) L35 [Wll] @ I N'DEX( SD: SBS7 1 .SD:SIS87,L33,SD:SFS119) K 35 (Wll) £1NDEX(SD:S BS71 ,SD:SIS87,M33,SD:SFS119) 136 [ W17] •OXTC J 36 ( F4) [Wll] @1F(J28="press ,3*SG:$DS82*SG:SDS84*J32,1000000"J34"(l-J14)/( $G:SDS85"(1+(1/SC:SFS36))),3'SG:SDS83*SG:SDS84"J32"1000000*J34*(1-J14)/(2*SG:S DS85 *(1*(1/SC:SFS36)))) J : K 36: (F4) [Wll] @IF(K28="press.",3'SG:SDS82*SG:SDS84*K32•1000000*K34•(1-K14)/( 2*SG : SDS85*(l+( 1/SC:SFS36)) ) , 3 * SG: SDS 83" SG: SDS 84 *K32" 1000000'K34 *(1-K 14) / ( 2* SG: S DS85"(1•(1/SC:SFS36)))) J :L36 : (F4) (Wll] @IF(L28»"press. ",3*SG:SDS82*SG:SDS84"L32*1000000*L34*(l-l.l4)/( 2"S3:SD$85*(1+(1/SC:SFS36))),3•SG:SDS83•SG:SDS84•L32•1000000*L34•(I-L14)/(2•SG:S DS 6 5 •( l +(l/SC:SFS36) ))) J : M 36 : (Ft) [Wll] 0IF(M28»"press.".3•SG:SDS82*SG:SDS84*M32•1000000*M34•(1-M14)/( 2*SG:SDS05"(1+(1/SC:SFS36))),3 *SG:SDS83•SG:SDS84*H32* 1000000*M34*(1-H14)/(2 *SG:S DS85-(!+(1/SC:SFS36)))) J: 137: [W17] 'FTC J:J37: (F4) [Wll] £IF(J28="press.",3*SG:SDS76*SG:SDS78•J32•1000000*J35"(1-J14)/( 2*SG:SDS79*(1+SC:SFS36)),3*SG:SDS77*SG:SDS78"J32*1000000*J35*(1-J14)/(2'SG:SDS79 *(1 + SC : SFS36))) J : K 37 : (F4) [Wll] 0IF(K28«"press.",3*SG:SDS76'SG:SDS78*K32*1000000*K35*(1-K14)/( 2*SG:SDS79*(1+SC:SFS36)),3*SG:SD$77*SG:$DS78*K32*1000000"K35*(1-K14)/(2*SG:$DS79 *(1+SC:SFS36))) J:L37: (F4) [Wll] ^IF(L28«"press.",3*SG:SDS76*SG:SDS78«L32*1000000"L35*(1-L14)/( 2*SG:SDS79*(1+SC:$FS36)).3•SG:SDS77«SG:SDS78»L32•1000000«L35•(1-L14)/(2•SG:SDS79 *(1+SC:SFS36))) J: M37 : (F4) [Wll] (IIF(H28* "press. " ,3 *SG : SDS76*SG :SDS78*M32* 1000000*M35*(1-M14 )/( 2*SG:SDS79'(1+SC:SFS36)),3*SG:$DS77*SG:SDS78*M32*1000000*M35*(1-M14)/(2*SG:SDS79 *(1 + SC:SFS36))) J:138: [W17] ' FSC .1: J38: (FA) [Wll) gIF(J28« 'press.SG:SDS29"SC:SFS 37/500,SG: SDS30"flEXP( 0.000759" $C:SFS37)) J:K38: (FA) [WLL] @IF(K28="press.",SG:SD$29*SC:SF$37/500,SG:$D$30*@EXP(0.000759* SC:SFS37)) J:L38: (FA) [WLL] flIF(L2B» "press. " , SG:$DS29"SC: SFS 37/500 ,$G:$DS30"@EXP(0.00D759* SC:SFS37)) J:M38: (FA) (WLL] f>IF(M28« "press.",SG:SDS29»SC:SFS37/500,SG:$D$30"§EXP(0.000759" SC:SFS37)) J HAO [W3] ' INTERMEDIATE MASS CALCS J I A 1 [W17] 'COEFSUM J JA1 (FA) (Wll] +SG:$DS35+SG:SDSA8+J37+J36+SG:SDS59+SG:SDS60+J38+SG:SDS22 J KA1 (FA) [WLL] +SG:SDS35+SG:SDSA8+K37+K36+SG:SDS59+SG:SDS60+R36+SG:SDS22 J LAI (FA) (Wll) +SG:SDS35+SG:$DSA8+L37+L36+SG:SD$59+SG: SDS60+L38+SG;SDS22 J MAI (FA) [Wll] +SG:SDS35+SG:SDSA8+M37+H36+SG:$DS59+SG:SDS60+M38+SG:SDS22 J IA2 [W1 7 1 'COEFTL J JA2 (FA) |W11] piF(M7»"YES"«AND«M6»l,SG:SDS1A+JA1,JA1) J KA 2 (FA) [Wll] §IF(M7="YES"tfAND«M6«2,SG:SDSlA+KAl,KAl) J LA2 (FA) (Wll] {)IF(M7»"YES"«AND''M6«3,SG:SDS1A + LA1,LA1) J MA 2 (FA) [Wll] piF(M7*"VES"wAND'/M6sA,SG:SDSlA + MAl,MAl) J IA3 [W17 J 1 COEFTLH J JA 3 (FA) (Wll] gIF(J7= "N0",0.@IF(M7 = "YES"«AND"M6 = l,SG.' SDS1A+SG:SDS35+SG:SDS60 ,50: SDS35+ SG :SDS60)) J : K A 3: (FA) [Wll] 0IF(J7="NO",0,§IF(H7="YES"HAND»H6«2,SG:SDS1A+SG:SDS35+SG:SDS60 ,SG:SDS35+SG :SDS60)) J : LA 3: (FA) (Wll) gIF(J7 = "N0",0.£IF(M7 = "YES"«AND»0,(J56+£ABS(J57)*0.5)"(l/3),(-!)•(($A BS(J56+^ABS(J57)*0.5))"(l/3))) J.-K58: < F J 160 (W17) 'MF (kg) J J60 (F4) ( Wll) +J58+J59-J51/3 J K60 (F4) ( WllJ +K58+K59-K51/3 J L60 (F4) ( Kill +L50+L59-L51/3 J M60 (F4) ( WH) +M58+M39-H51/3 J H62 [W5] ' HASS SUMMARY J 163 (W17) '«A (kg) J J63 (F2> [ WllJ @IF(M7« " Yes "0AND//M6" 1 ,$G:SD$14*(J86-J6 M73 F2) [Wll] +$G:SD$ P16 [W17 'SFB [W/B2] Q16 [Wll +SG:SDS61"(SG:SDS13/Q15)"2 R16 [Wll +SG:SDS61'(SG:SDS15/R15)"2 S16 [Wll +SG:SDS61"(SG:$D$15/S15)*2 T16 [Wll +SG:SDS61«($G:SDS15/T15)"2 018 (W5) ' REFRIGERATION MASS PRELIMINARY CALCULATIONS P19 [W17 ' Oxidizer S19 (Wll ' Fue 1 P20 [ W17 1HOX [W/n2] Q20 (FA) [Wll] +SG:SD$58'$G:$D$21*$G:$D$65'4 S 20 [Wll •HF T20 (FA) [Wll] +SG:SDS58,SG:SDS20'SG:SDS63*A P21 (W17 'ROXT [ra/(kg)"1/3) Q 2 1 (FA) [Wll] (0.75/(SG:SDS19"SG:SDS50"§IPl))"(l/3) S 2 1 [Wll ' RFT T21 (FA) (Wll] (0.75/(SG: SDS17»$G: $D$A9«(!PI)) *(1/3) P22 [W17 ' QIN [W/( kg)' 2/3] Q22 (FA) [Wll] (Q16"(l-SG:SDS5A)"SG:$DS2A)'pPI"Q21-2 S22 (W1 1 •QIN T22 (FA) I W11) (T16"(l-SG:SDS3A)*$G:SDS23)*ePI'T21-2 P 2 3 [W17 QUL'T [ " " ] Q23 (FA) [tall) CQ20-SG:SDS32)* A'(IPI* Q21"2 S 2 3 [Wll ' QOUT T23 (FA) [Wll] (T20"$G:SDS31)"A*gPl"T21"2 P2A [ W 17 'QCOOL [ " ' ) Q2A (FA) (Will +SG:SDS50"(Q22-Q23) S2A [Wll 'QCOOL T2A (FA) (tall] +$G:SD$A9,(T22-T23) P 2 5 [ W 17 '"RPRE [kg*1/3] Q25 (FA) (Wll] (SG:SDSAA,Q2A/$G:$DS52) S25 [Wll 'MRPRE T25 (FA) [Wll] (SG:SDSA3'T2A/SG:SDS51) 027 I W5] ROCKET DESIGN DETAILS (stage dependence) P 20 ( W17 'FEED SVSTEH USED Q28 [Wll "press. R 2 0 iwn "press. S 28 [Wll 'press. T20 IH 11 "press . P29 [W17 •THRDES [kg] Q29 (F0) [Wll] +SG:SDS66"Q86 R29 (FO) [Wll] +SG:$DS66"R06 S29 (FO) (Wll) +SG:SDS66"S86 T29 (FO) [Wll] + SG:SD566"T86 P30 (F2) [W17) 'TIMEB [sec) Q30 (F2) [Wll] +Q79'Q13/Q29 R30 (F2) [Wll] gIF(R29«0,0iR79*R13/R29) S 30 (F2) [Wll] eiF(S29*0,0,S79«S13/S29) T30 (F2) [Wll] £IF(T29»0,0.T79*T13/T29) P31 [ W 17] 'OPT. FEED SYSTEM Q31 [Wll] glF(Q30<»(8130.1'Q29--0.75867)," press."," pump") R31 [Wll) g IF( R29*0." n/a",§IFCR30<«(8130.l*R29--0.75867)," press. pump" )) J :S31 : [Wll] @1F ( S 29 «0 , " n/a",#IF(S30<«(8130.1'S29--0.75867)," press. pump" )) J :T31: [Wll] gIF(T29*0," n/a",eiF(T30<»(8130.1 *T29*-0.75867)," press. pump" )) J:P32: [W17] 'PT (HPaJ J:Q32: (F4) [Wll] glF(Q28« •PUMP",50*0.101323/14.696,1.5*SC:$FS37*0.101323/14.696 ) J:R32 : (F4) [WU] 0IF(R20«"PUMP",50*0.101323/14.696,1.3*SC:SFS37*0.101325/14.696 ) J:S32: (F4) [Wll] 0IF(S28«"PUMP",50*0.101325/14.696,1.3*$C:$FS37*0.101323/14,696 ) J:T32: (F4) [Wll] 0IF(T28="PUMP",30*0.101325/14.696,1.3"$C:$FS37"0.101323/14.696 ) J:P33: (W17] 'PINDEX J:Q33: (F0) [Wll J piF(Q32>«SD:SIS71,7,gIF(Q32>»SD:$HS71,6,gIF(Q32>=SD:SG$71,5,pi F(Q32>=SD:SFS71,4,gIF(Q32>-SD:SES71,3,SD:SJS61))))) J:R33: (FO) [Wll] piF(R32>"SD:SIS71,7,01F(R32>«SD:SHS71,6,piF(R32>•SD:SGS71,5,01 F(R32>=SD:SFS71,4,gIF(R32>-SD:SES71,3,$D:SJS61))))) J: S 3 3: (FO) [Wll] gIF(S32> = SD:SIS71,7,gIF(S32>=>SD:SHS71,6,pIF(S32> = SD:SGS71 ,5,01 F(S32>=SD:SFS71,4,0IF(S32>"SD:SES71,3,SD:SJS61))))) J:T33: (FO) [Wll] 01F(T32>«SD:SIS71,7,01F(T32>=SD:SHS71,6,piF(T32> = SD:SGS71,5,01 FCT32>»SD:$FS71,4.gIFCT32>"SD:SES71,3,SD:SJS61))))) J:P34: IW17] 'VOX |n3/kg] J:Q34: [Wll] 0INDEX(SD:SBS71, •SD:SIS87,Q33,SD:SFS120) J :R34: (Wll) gINDEX(SD:SBS71. .SD:SIS87.R33,SD:SFS120) J:S34: [Wll] 0INDEX(S D:SBS71. .SD:SIS07,S33,SD:SF512O) J:T34: [Wll] 01 S'DEX( SD:SBS71. . SD:SIS87,T33,SD:SFS120) J:P35: [W17] •VF [m3/kg] J:Q35: (Wll) gINDEX(SD:SBS71..SD:SIS87.Q33,SD:SFS119) J:R35: [Wll] PINDEX(S D:SBS71..SD:SIS87.R33.SD:SFS119) J:S35: [Wll] PINDEX(SD:SBS71. -SD:SISB7,S33,SD:SFS119) J.-T35: [Wll] glS'DEXf SD: SBS71 ..SD:SIS67,T33,SD:SFS119) J:P36: IW171 'OXTC J:Q36: (Ft) [Will 01F(Q283"press.",3*SG:SDS82"SG:SDS64*Q32•1000000"Q34•(1-Q14 )/( 2*SG:SDS85"(1 +(1/SC:SFS36))),3 *SG:SDS83«SG:SDS84*Q32* 1000000"Q34•(1-Q14)/(2'SG:S DS85 "(1 •»( 1 /SC:SFS36)) ) ) J:R36: (F4) [Wll] £IF(R28="press3*SG:SDS82 *SG:SDS64*R32"1000000 *R34"(1-R14)/( 2,SG:SDS85"(1+(1/SC:SFS36))),3*SG:SDS83*SG:SDS84*R32*1000000,R34*(1-R14)/(2*SG:S DS85 • ( 1M 1 /SC: SFS36 ) ) ) ) J:S36: (Ft) [Wll] 01F(S28-"press3•SG:SDS82«SG:SDS84•S32*1000000•S34"(1-S14)/( 2"SG:SDS85,(1+(1/SC:SFS36))),3"SG:$DS83*SG:SDS84*S32*1000000*S34'(1~S14)/(2"SG:S DS85'(It(1/SC:SFS36)) ) ) J:T36: (F4) [Wll] 01F(T28="press.",3*SG:SDS82•SG:SDS84«T32* 1000000*T34•(1-T14)/( 2"SG:SBS85,(l+(l/SC:SFS36))),3*SGiSDS83*SG:SDSB4«T32"1000000*T34'(l-T14)/(2"SG:S DS85*(1+(1/SC:SFS36)))) J : P37: [W17] 'FTC J:Q37: (F4) [Wll] piF(Q28«"press.",3*SG:SDS76*SG:SDS78•Q32*1000000•Q35•(1-Q14)/( 2*SG:SDS79"(1+SC:SFS36)),3*SG:SDS77•SG:SDS78*Q32*1000000*Q35*(1-Q14)/(2 *SG:SDS79 "(1+SC:SFS36))) J:R37: (F4) [Wll] pIF(R28«"press.".3•SG:SDS76•SG:SDS78*R32•1000000•R35*(1-R14)/( 2*SG:SD$79*(1+SC:SFS36)),3"SG:SDS77*SG:SDS78"R32"1000000*R35*(1-R14)/(2*$G:SDS79 "(1+SC:SFS36))) J:S37: (Ft) [Wll] 0IF(S28»"press3*SG:SDS76*SG:SDS78*S32* 1000000'S35"(1-S14)/( 2*SG:SDS79*(1+SC:$FS36)),3"SG:SDS77"SG:SDS78*S32*1000000*S35*(1-S14)/(2*SG:SDS79 •( 1 + SC :SFS36 ))) J: T37: (F4) (Wll] 0IF(T28-"press . " , 3 * SG :SDS76* SG: SDS78*T32 * 1000000*T33*( 1-T14 ) / ( 2*SG:SDS79*(1+SC:SFS36)),3*SG:SDS77,SG:$D$78*T32*1000000"T35"(1-T14)/(2*SG:SDS79 '(1+SC:SFS36))) J:P3B: (W17] ' FSC J:Q38: (F*t) (Wll) gIF(Q28="pressSG:SDS29*SC:SFS37/500,SG:SDS30*fEXP(0.000759" SC:SFS37)) J: R3B: (FA) [Wll] plF(R28»"press.",SG:SDS29"SC:SFS37/500,SG:$DS30,flEXP(0.000759" SC:SFS37)) J:S3B: (FA) (Wll) gIF(S28»"press.",SG:SDS29"SC:SFS37/500,SG:SDS30*gEXP(0.000759* SC:SFS37)) J:T39: (FA) (Wll) 0IF(T28«"pressSG:SDS29*SC:SFS37/500,$G:SD$30*eEXP(0.000759* SC :SFS37)) J: 0A0: (W5] ' INTERMEDIATE MASS CALCS J:PA 1: [W17] 'COEFSUM JiQAl: (FA) (Wll) +$G:SDS35+SG:$D$A8+Q37+Q36+SG;SDS59+SG:$DS6O+Q30+SG:$DS22 JiRAl: (FA) (Wll) +SG:SDS35+SG:SDSA8+R37+R36+SG:SDS59+SG:SDS60+R38+SG:SDS22 J:SA1 : (FA) (Wll) +SG:SDS35+ SG:SDSA8+S37+S36+SG:SDS59+SG:SDS60+S38+SG:SDS22 J:TA 1 : (FA) (Wll) +SG:SDS35+SG:SDSA8+T37+T36+SG:SDS59+SG:SDS60+T38+SG:SDS22 J:PA2: (W17) 'COEFTL J:QA2: (FA) (Wll) gIF(T7«"YES"«AND«T6»1,SG:SDS1A+QA1,QA1) J:RA2: (FA) (Wll] 0IF(T7=»YES"WAND#T6«2,SG:SDS1A+RA1,RA1) J :S A 2: (FA) |W11] 0IF(T7= " YES "«AND«T6"=3 , SG:SDS 1A + SA 1.SA 1) J : T A 2: (FA) (Wll) piF(T7«"YES"«AND«T6= A,SG:SDS1A + TA1,TA1) J : PA 3: (W17] 'COEFTLM J:QA 3: (FA) (Will 0IF(Q7="NO",0,£IF(T7="YES"»AND«T6=1,SG:SDS1A+SG:SDS35+SG:SDS60 ,SG:SDS35+SG:SDS60)) J:RA3: (FA) (Wll] 0IF(Q7="NO",0,0IF(T7="YES"«AND«T6=2,SG:SDS1A+SG:SDS35+SG:SDS60 ,SG:SDS35+SG:SDS60)) J :S A 3: (FA) [Wll] piF(Q7« "NO",0 ,0IF (T7« " YES "«AND/'T6 = 3 ,SG : SDS 1 A + SG :SDS35 + SG:SDS60 ,SG:$DS35+SG:SDS60)) J :TA 3: (FA) [Will £IF(Q7 = "NO",0,01F(T7»"YES"0AND0T6* A,SG:SDS1A + SG:SDS35 + SG : SDS60 ,SG:SDS35+SG:SDS60)) J PA A [W17] 'NUMER J QA A (FA) (Wll) +Q1A-QA2 J RAA (FA) (Wll] +R1A-RA2 J SAA (FA) [Wll| +S1A-SA2 J TA A (FA) [Wll] +T1A-TA2 J PA 5 [W17] 'DENOM J QA5 (FA 1 [W11J +SG:SDS56*( 1 -Q1A) J RA 5 (FA) (Wll] +SG:SDS 56" (1 -R1A) J SA 5 (FA) (Wll] +SG:SDS56* (1 -SIA) J TA 5 (FA) (Wll] +SG:SDS 56' ( 1 -T1A) J PA6 (W17] 'VALUE J QA6 (FA) [Wll] 100*QAA/QA 5 J RA6 (FA) (Wll) @IF(RA5*0,0, 100*RAA/RA5) J SA6 (FA) [Wll] EIF(SA5=0,0,100*SAA/SA5) J TA6 (FA) (Wll] EIF(TA5=0,0,100*TAA/TA5) J PA7 ( W17] •VAL2 J QA7 (FA) (Wll] 1-QA2 J RA7 (FA) [Wll] 1-RA2 J SA7 (FA) [Wll] 1-SA2 J TA 7 (FA) [Wll] 1-TA2 J PA 8 (W17) ' V2ML J QA8 (F2) [Wll] +QA7*Q11 J RA 8 (F2) (Wll] +RA7*R11 J SAB (F2) [Wll] +SA7«S11 J TA 8 (F2) [Wll] +TA7-T11 J:050: (W5) ' SOLVING CUBIC FOR MF J: P51 : [Vr 17] 1 A1 J:Q51: (S3) [Wll] £IF(Q8= "No"»'AND«T8»"No" ,-(3"Q J P60 [W17] 'MF (kg) J Q60 (FA) Wll] +Q58+Q59-Q51/3 J R60 (F<0 Wll] +R58+R59-R51/3 J S60 (FA) Wll] +S58+S59-S51/3 J T60 (FA ) Wll] +T56+T59-T51/3 J 062 [W5J MASS SUMMARY J P63 [W17] 'MA [kg] J Q63 (F2) *11] 0IF(T7»"Yes"«AND#T6«l,$G:SD$l J T73 (F2) [Kill •SG:$D$47 • (T86-T6'i) J P74 [W17] •MFT (kg) J Q74 (F2) [Wll] +Q37•(Q86 -Q64) J R74 (F2) (¥11] +R37•(R86-R64) J S74 (F2) [Wll] +S37"(S86-S64) J T7*. (F2) [Wll] +T37•(T86 -T64) J P75 [K17] 'HOXT (kg) J Q75 (F2) [Wll] +Q36'(Q86-Q6<0 J R75 (F2) [Wll] +R36*(R86-R64) J S75 (F2) [Wll] +S36•(S86-S64) J T75 (F2} [Wll] +T36"(T86 -T6A) J P77 (K17) •MF [kg] J Q77 (F2) [Wll] +Q60 J R77 (F2) [Wll] + R60 J S77 (F2) [Wll] +S60 J T77 C F2 J [Wll] +T60 J P70 [W17J ' MOX (kg) J Q70 (F 2 ) [Wll] +Q77 * $C:$FS36 J R78 (F2) [Wll] +R77•SC:SFS36 J S78 (F2) [Wll] +S77•SC:SFS36 J T78 (F2) (Wll] +T77•SC:SFS36 J P79 [ W17 J •HP (kg) J Q79 (F2) [Wll] +Q77+Q78 J R79 (F 2) [Wit] +R77+R78 J S79 (F2) [Wll] +S77+S78 J T79 (F2) [Wll] +T77+T78 J P 81 [ W17 J 'NUMMOD J P82 I W17] 'MMODTL [kg] J P83 [ W1 7 ] 'MODTHRTL (kg] J P 8 5 I W17] ' Ml [kg] J Q85 (F2) [Wll] PSUM(Q63. • Q75) J R05 (F2) [Wll] gSUH(R6 3. R75) J S85 lF2) [Wll] pSUM(S63.. S75) J T85 CF2) [Wll] pSL'M(T63 . T75) J P8b IW1 7 ] ' PO [kg] J Q06 (F2) [Wll] +Q79/(1-Q14) J R66 (F2) [Wll] glFCRll'l ,0,R79/(1- J S8fi (F2) [Wll] ?IF(SH = 1 ,0,S79/(1- J T86 (F2) (Wll) gIF(TH«l i0,T79/(1- J XI: 1W11 +SC:SAS22 J V2: |W5| •SB:SBS 29 J X2: (Wll +SB:SCS29 J V 3: [V5| \" J W3: I W 17 \» J X3: [Wll J Y3 [Wll \» J Z3: [Wll \« J AA3: [Wi ) \" J WA: (W17 'TOTAL DELTA V J XA: [Wll +SB:SGS29+SB:SGS30 J YA: [Wll TOTAL tt OF STAGES J AAA : [WI ] +SB:SES 31 J W5: [ W17 •MODULAR ENGINES J X5: [Wll + SB:SES33 J Y5: [Wll AEROBRAKING J AA5: (WI ) +SB:SES32 J W6: [ W 1 7 'OX. REFRIG. J X6: [Wll + SC:SES 26 J Y6: [WI 1 FUEL REFRIG. J AA6 : in ] +SC:S ES 2 5 J W7 : [ W 1 7 •OX. SURF. REF. J X7: [Wll •SC:SFS35 J Y7: [Wll FUEL SURF. REF. J AA7: |W1 ) +SC:SFS3A J W8: [ W 1 7 'OX. ISRU J X8: [Wll + SC:SGS 35 J Y8: [Wll FUEL ISRU J AA8 : [Wi I +SC:SGS 3A J XI0: |W1 ) "STAGE I J Y10: |W1 ) -STAGE II J Z10: [Wi ) "STAGE III J AA10: |W 1] 'STAGE IV J Wll : |W1 ) 'ML [kg] J Xll; (F2 (Wll) £IF(AAA=1,Q86+$G:SDS26'Q77+SG:$DS26,Q78+C93+F9A+F95+F103+F111- F119-F127 , 1F(AAA >1,Y06 + SG:SDS25,Y77+ SG:SDS27'Y70,O)) J VI1: ( F 2 [Wll] gIF(AAA«2.Q86+SG:SDS26•Q77+SG:SDS28*Q78+C93+F9A+F95+F103+F111- F119-F127, IF(AAA>2iZ86+SG:$DS25"Z77+SG:SDS27*Z78.0)) J Zll : (F2 [Wl 1] gIF(AAA=3,Q86+SG:SDS26'Q77+ SG:SDS26"Q70+C93+F9A + F95 + F103+ F111 - F119-F127,^IF(AA«>3,AA86+SG:SDS25,AA77+ SGiSDS27"AA78,0)) J A A11 : (F ) [Wll] £IF(AAA»A,Q86+$G:SDS26*Q77+$G:SDS28*Q78+C93+F9A+F95+F103+F111 -F119-F1 27 0) J W12: [WI ) 'DV [n/s) J X12: [WI ) +XA/AAA J Y12: [Wi ) gIF(AAA>«2,XA/AAA,0) J Z12: [Wi ) £IF(AAA>*3,XA/AAA,0) J AA12: [W I) gIF(AAA>»A,XA/AAA,0) J W13: [WI ) 'ISPUSED {S) J X1 3: [WI ) +SG:SDS 38 J Y13: [WI j +SG:SDS38 J Z13: [Wi 1 +SG:SDS 38 J AA13: [W 1) +SG:SDS38 J W1A: [Wl ) 'MRATIO J X1A : (FA [Wll] §EXP(-X12/(X13*SG:SDS37)) J Y1A: (FA [Wll] PEXP(-Y12/(Y13*SG:SDS37)) J Z1A : (FA [Wll] gEXP(-Z12/(Z13'SG:SDS37)) J AA1A : (F ) IVill] £EXP(-AA12/(AA13"SG:$DS37)) J W15: [*17 •RBS [AU] J X15: (W11 +SB:SCS6 J ¥15: [ W11 +SB:$CS6 J Z15: [*11 +SB:SCS6 J AA15 [VI| +SB:SCS6 J W16: (W17 'SFB {W/b2} J X16: [Wll •SG:SDS61*(SG:$D$15/X15)"2 J Y16: (W 11 +SG:SDS61"(SG:SDS15/Y15)"2 J Z16: [WU +SG:SDS61"($G:$DS15/Z15)*2 J AA16 IW1 ] +SG:SDS61,(SG:SDS15/AA15)"2 J V16: 1*5] • REFRIGERATION MASS PRELIMINARY CALCULATIONS J W19: [W17 ' Oxidizer J Z19: [Wll Fue 1 J W20: [W17 'HOX [W/D2] J X20: (FA) [Wll] +SG:SDS58"SG:SDS21*$G:SDS6A"A J Z20: IW 1 1 • HF J AA20 ( F A [Wll] +SG:SDS58*SG:SDS20'$G:SDS62"A .1 v» 21: [ W 17 •ROXT [m/(kg)* 1/3) J X21 : (FA) [Wll] C0.75/(SG:SDS18*SG:SDS50-ePI))-(l/3) J Z21: [Wll ' RFT J AA21 (FA [Wll] (0.75/(SG:SDS16*SG:SDS49a@PI))*(l/3) J W22: 1 W 1 7 •QIN [W/(kg)"2/3] J X22: ( FA ) |W11] (X16'(1-SG:SDS5A)«SG:$DS2A)«0PI"X21"2 J Z22: [Wll ' QIN J AA22 (FA [Wll] (AA16"(1"SG:SDS3A)"SG:SDS23)•gPI*AA21*2 J W 2 3: [ W 17 ' QOl'T [ " » ] J X23: (FA ) [Wll] (X20"SG:SDS32)*A,§PI«X21_2 J Z23: [Wll ' QOUT J AA23 (FA [Wll] (AA20*SG:SDS31)"A'@PI*AA21"2 J W2A: [ W 17 'QCOOL [ " " ] J X 24 : (FA) [Wll] +SG:SD$50"(X22-X23) J Z2A: [Wll •QCOOL J AA 2A (FA [Wll] tSG:SDSA9"('AA22-AA23) J W25: [W17 'MRPRE [kg * 1/3] J X 2 5: (FA) [Wll] ($G:$DSAA*X24/$G:$D$52) J Z 2 5: [Wll 'MRPRE J AA25 (FA IW11 ] (SG:SDSA3"AA2A/SG:SDS51) J V27 : 1W5 ] ' ROCKET DESIGN DETAILS (stage dependence) J to 28 : IW 17 'FEED SYSTEM USED J X20: [Wll " pump J V28: [Wll "pump J Z 28 : [Wll "punp J A A 20 [W1 ] "pump J W29: [ W17 'THRDES [kg] J X29: (F0) [Wll] +SG:SD$66*X66 J V 29: (FO) [Wll] +SG:SDS66*Y06 J Z29: (FO) [Wll] +$G: SD$66*Z86 J AA29 (FO) [Wll] +SG:$D$66*AA86 J W30: (F2) [W17] 'TIMEB [sec) J X 3D: (F2) [Wll] + X79 * X13/X29 J Y 30: (F2) (Wll) piF(Y29»0,0,Y79"Y13/Y29) J Z3U: (F2) [Wll] piF(Z29«0,0,Z79"Z13/Z29) J AA30 (F2 ) [Wll] @IF(AA29«0,0,AA79*AA13/AA29) J: W31 : |W17] 'OPT. FEED SYSTEM J:X31: (Wll) 0IF(X3O<»(013O.l*X29*-0.75867)," press."," punp") J!Y31: [Wll) 0IF(Y29-O," n/a",0IF(Y3O<-(8130.1»Y29'-0.75867)," press.", " punp")) J:Z31: [Wll] 0IF(Z29=O," n/a",0IF(Z3O<«(8130.l*Z29--0.75867)," press.", pump")) J:AA31: (Wll) 0IF(AA29«O," n/a",0IF(AA3O<-(813O.l*AA29--0.75867)," , pres s."," pu«p")) J:W32: [W17] 'PT [MPa) J:X32: (F4) [Wll] 0IF(X28-"PUMP",50*0.101325/14.696,1.5•SC:SFS28»0.101325/14.696 ) J:Y32 : (F4) [Wll] 0IF(Y28-"PUMP" , 50*0.101325/14. 696 ,1.5 * SC: $FS28*0 .101325/14 . 696 ) J :Z32 : (F4) [Wll] 0IF(Z2S» "PUMP" , 50*0.101325/14. 696,1. 5«SC :SFS28*0 . 101 325/14. 696 ) J:AA32: (F4) [Wll] 0IF(AA28S"PUMP",50*0.101325/14.696,1.5*SC:SFS28*0.101325/14.6 96) J:W33: [W17] 'PINDEX J : X33: (FO) (Wll) 01F(X32> = SD :S 1 S7 1 .7 ,0IF (X32> = SD:SHS71 ,6 ,0IF (X32>-SD: SGS71,5 ,01 F(X32>«SD;SFS71,4,(UF(X32>«SD:SE$7L,3,SD:$JS61))))) J : Y33: (FO) [Wll] 01F(Y32> = SD:S1S71,7,0IF(Y32>»SD:SHS71,6,0IF(Y32>»SD;SGS71,5,01 F(Y32>=SD:SFS71,4,piF(Y32>=SD:$ES71,3,SD:SJS61))))) J:Z33: (FO) (Wll) 01F(Z32>=SD:S1$71,7I0IF(Z32>«$D:SHS71,6,0IF(Z32>«SD:SGS71,5,0I F(Z32>»SD;SFS71,4,01F(Z32>=SD:SES71,3,SD:SJ$61))))) J:AA33: (FO) [Wll] 0IF(AA32>"SD:SIS71,7,0IF j AD16 [ W17] 'SFB (W/n2J j AE16 [Wll] +SG:SDS61*(SG:SDS15/AE1S)"2 j AF16 [Wll] +SG:SDS61*(SG:$DS15/AF15)',2 j AG 16 [Wll] +SG:SDS61'(SG:SDS15/AG15)"2 j AH 16 [Wll] +SG:$DS61*(SG:SDS15/AU15)-2 j AC18 [W5] ' REFRIGERATION MASS PRELIMINARY CALCULATIONS j AD19 [W17] ' Oxidizer j AG 19 [Wll] 1 Fuel j AD20 [W17] 'HOX [W/n2] j AE20 (F^<) [Wll] +SG:SDS58,SG:SDS21"$G:$DS6«-« j AG20 [Wll] 'HF j AH20 (F4) [Wll] +SG:$D$58*$G:$D$20*$G:$D$62"4 j AD21 IW17] ' ROXT [ni/(kg) * 1/3] j AE21 (F4) [Wll] (0.75/(SG:SDS18,SG:SDSS0*pPI))*(l/3) j AG21 [Wll] 'RFT j AH21 J:AD38: [W17 ] 'FSC J:AE38: (FA) (Wll) pIF(AE28" •press.SG:$DS29•SC:SFS28/500,SG:SD$30»pEXP(0.00075 9"SC:SFS28)) J:AF38: (F4 3 (WU) piF(AF28" 'press. ',SG:SDS29,SC:SFS28/500,$G:$DS30*pEXP(0.00075 9«SC: SFS20)) J:AG30: (FA) [Wll] piF(AG28*"press. ',SG:$DS29'SC:$F$28/500,$G:$D$30'PEXP(0.00075 9"SC:SFS28)) :AH30: (FA) [Wll) PIF(AH28="press.",SG:SDS29"SC:SFS28/500,SG:$DS30*pEXP(0.00075 "SC:SFS28 )) :ACA0: [W5 ] INTERMEDIATE MASS CALCS :ADA1: [W17) 'COEFSUM :AEA1: (FA) [Wll] +SG:SDS35+SG:$D$A7+AE37+AE36+$G:$D$59+$G:$DS60+AE38+SG:SDS22 :AFAl: (FA) [Wllj +SG:SDS35+SG:SDSA7+AF37+AF36+SG:SD$59+SG:SDS60+AF38+SG:$DS22 :AGA 1: (FA) [Wll] +SG:SDS35+SG:SDSA7+AG37+AG36+SG:SDS59+SG:SDS60+AG38+SG:SDS22 :AHA1: (FA) [Wll] +SG:SDS35+SG:SDSA7+AH37+AH36+SG:SDS59+SG:SDS60+AH38+SG:SDS22 : ADA 2: [W17] 'COEFTL :AEA2: (FA) (Wll) PIF(AH7 = "YES"'/AND«AH6»1,$G:SDS1A + AEA1,AEA1) : A F ^ 2: (FA) [Wll] piF(AH7«"YES"«AND»AH6=2,SG:SDSlA+AFAl,AFA1) : AG*" 2: (FA) [Wll) piF(AH7»"YES"wANDWAH6=3,SG:$DSlA+AGAl,AGAl) : A H 4 2: (FA) [Wll] piF(AH7»"YES"WAND«AH6=A,SG:SDS1A+AHA1,AHA 1) : ADA 3: [W17) ' COEFTLM :AEA3: (FA ) [Wll] piF(AE7="N0" ,0 , pi F(AH7 = " YES " It AND//AH6" 1 , SG i SDS1A + S G :SDS 35+ SG: S DS60,SG:SDS35+SG:SDS60)) J : AF A 3: (FA) [Wll] piF(AE7= ' NO",0.pIF(AH7»"YES"«AND«AH6=2,SG:SDS1A+SG:SDS35+SG:S DS60,SG:SDS35+SG:SDS60)) J : AGA 3: (FA) (Wll) piF(AE7=' NO" ,o.pIF(AH7«"YES"«ANDwAH6=3,SG:SDS1A+SG:SDS35+SG:S DS60,SG:SDS35+SG:SDS60)) J : AHA 3: (FA) [Wll] piF(AE7-' NO' ,0,piF(AH7»"YES"«ANDtfAH6=A,SG:SDS1A+SG:SDS35+SG:S DS60 , SG:SDS35 + SG:SDS60)) ADAA [W 17] 'NUMER AEAA (FA) (Wll) +AE1A-AEA2 AF A A (FA) (Wll) +AF1A-AFA2 AG A A (FA)| W11) +AG1A-AGA2 . AHAA (FA) (Wll) +AH1A-AHA2 ADA 5 [W17] 'DENOM AEA5 (FA) [Wll] +SG:SDS551(1-AE1A) AF A 5 (FA) [Wll] +SG:SDS55*(1-AF1A) AGA 5 (FA) [Wll] +SG:SDS55*(1-AG1A) AHA 5 (FA) [Wll] +SG:SDS55"(1"AH1A) ADA6 [WI7] 'VALUE AEA6 (FA) [Wll] 100•AEAA/AEA5 AF A 6 (FA) [Wll] piF(AFA5=0,0,100'AF A A/AFA5) AG A 6 (FA) [Wllj piF(AGA5=0,0,100•AGAA/AGA5) AHA6 (FA) [Wll] PIF(AHA5=0,0,100"AHAA/AHA5) ADA7 [W17) 'VAL2 AEA7 (FA) [Wll] 1-AEA2 AFA7 (FA) [Wll] 1-AFA2 AGA7 (FA) (Wll] 1-AGA2 AHA7 (FA) [Wll] 1-AHA2 ADAB [ W17] 'V2ML AEA 8 (F2) [Wll] +AEA7• AE11 AF A 8 (F2) [Wll] +AFA7•AF11 AGAB (F2) [Wll) +AGA71AG 11 AHA8 (F2 ) (Wll) +AHA7" AH11 J:AC50: [W5] ' SOLVING CUBIC FOR MF J:AD51: {W17] 'A1 J: AE51: (S3) (Wll] (HF(AE8«"No"//AND//AH8""No".-(3*AEA8/AEA6),(JIF(AE8»"No",-(3"AEA 8/AEA6+(AH25/AEA6)-3),CIIF(AH8»"No",-(3*AEA8/AEA6+(($C:$FS27)-(2/3)'AE25/AEA6)-3) ,-(3*AEA8/AEA6+((AH25+($C:SFS27)-(2/3),AE25)/AEA6)-3)))) J: AF51 : (S3) (Wll] PIF( AFA6-0,0,?IF(AE8»"No"#AND#AH8«"NO" ,-(3*AFA8/AFA6) ,0IF( AE8 e "No",-( 3 *AFA8/AFA6+(AH25/AF46 )*3)I£IF(AH8""No"T-(3"AFA0/AF46+(($C:SFS27)"(2/3)* AE25/AFA6) *3) ,-(3,AF'I8/AF'T6+( (AH25+(SC:$FS27)"(2/3)•AE25)/AFA6)" 3))))) J: AG51: (S3) [Wll] £IF(AGA6»0 ,0 , £IF( AE8* "No "(*ANDf/AH8» "No" ,~(3*AGA8/AGA6),@IF (AE8 •"No" , -( 3" AG 48/AG 4 6+ ( AH25/AG46) *3),piF(AH8*"NoH,-(3*AGA8/AG46+(($C:$F$27)*(2/3)* AE25/AGA6)"3) ,-(3•AGA8/AGA6+((AH25+($C:SF$27)*(2/3)•AE25)/AGA6)* 3) )))) J: AH5 1 : (S3) [Wll] (IIF( AH46-0,0 ,^ IF (AE8»" No "(/AND//AH8-"No",-(3» AHA8/AHA6),$IF (AE8 = "No" ,-( 3*AHA8/AHA6+(AH25/AHA6)*3),0IF(AH6*"No" ,-(3*AHA8/AHA6+((SC:$FS27)*(2/3)• AE25/AHA6)-3),-(3•AHA8/AHA6+((AH25+(5C:SFS27)*(2/3)•AE25)/AHA6)* 3))))) J AD52 [W17] • A2 J AE52 (S3) (Wll] 3*(AEA8/AEA6)"2 J AF52 (S3) [Wll] @IF(AFA6»0i0t3'(AFA8/AFA6)"2) J AG52 (S3) [Wll] £IF(AGA6=0i0i3*(AGA8/AGA6)"2) J AH52 (S3) IWU] eiF(AHA6=0,0,3*(AHA8/AHA6)*2) J AD53 IW17] 'A3 J AE53 (S3) [Wll] -(AEA8/AEA6)*3 J AF53 (S3) [Wll] piF(AFA6=0,0,-(AFA8/AFA6)-3) J AG53 (S3) [Wll] @IF(AGA6=0,0,-(AGA8/AGA6)-3) J AH53 (S3) (Wll) gIF(AHA6»0,0,-(AHA8/AHA6)*3) J AD55 1 W17] 'Q J AE55 (S3) [Wll] (3*AE52-AE51"2)/9 J AF 55 (S3) [Wll] (3*AF52-AF51"2)/9 J AG55 (S3) [Wll] (3"AG52-AG51*2)/9 J AH55 (S3) [Wll] (3,AH52-AH51*2)/9 J AD56 [W17] •R J AE56 (S3) IWU] (9,AE51,AE52-27"AE53-2*AE51"3)/5A J AF56 (S3) IW11] (9"AF51"AF52-27*AF53-2*AF5l"3)/5A J AG 56 (S3) [WU] (9"AG51*AG3 2-27*AG53-2*AG51"3)/5A J AH56 (S3) [Wll] (9*AH51*AH52-27'AH53-2"AH51*3)/5A J AD57 [K17| •D J AE57 (S3) [Wll] •AE55"3+AE56 * 2 J AF57 (S3) [Will •AF53-3+AF56-2 J AG57 (S3) IWU] + AG55 * 3+AG56"2 J AH57 (S3) [WU] + AH55"3+AH56 * 2 J AD58 [W17] •s J AE58 (FA) [WU] §IF((AE56+§ABS(AE57) *0.3) >«0, (AE56+CIABS(AE57)*0.5)" (1/3), (-1 ) •((gABS(AE56+gABS(AE57)"0.5))*(l/3))) J:AF58: (FA) [Wll] § IF((AF56+0ABS(AF57)*0.5)>«0,(AF56+#ABS(AF57)*0.5)*(1/3),(-1 ) '((0ABS(AF56+0ABS(AF57)*0.5))"(1/3))) J:AG58: (FA) [Wll] gIF((AG56+AG57"0.5)>»0,(AG56+AG57*0.5)*(1/3),(-1)•((£ABS(AG36 +AG57-0.5))*(1/3))) J:AH58: (FA) [Wll] #IF((AH56+AH37"0.5)>«0,(AH56+AH37'0.5)"(1/3),(-1)•((0ABS(AH56 + AH57-0.5))*(1/3))) J:AD59: (W17] 'T J:AE59: (FA) [Wll] 0IF( (AES6-$ABS (AE57) *0. 3 ) <0 ,(-1) * ((#ABS (AE56-J1ABS (AE57) *0. 5 )) •(1/3)) , (AE56-0ABSUE57) "0.5)" (1/3) ) J:AF59: (FA) (Wll) glF((AF56-^ABS(AF57)*0.5)<0,(-1)•((gABS(AF56-gABS(AF57)*0.5)) "(1/3)),(AF56-$ABS(AF57)*0.5)"(1/3)) J:AG59: (FA) (Wll] F((AG36-AG37*0.5)<0,(-1)*((£ABS(AG56-AG57*0.5))*(1/3)),(AGS 6-AG57-0.5)*(1/3)) J:AH59: (FA) [Wll] §1F((AH36-AH57"0.5)<0,(-1)•((gABS(AH36-AH57-0.5))*(1/3)),(AH3 6-AH57-0.5)*(1/3)) AD60 [ W17] *MF [kg] AE60 (F2) (Wll] +AE58+AE59-AE51/3 AF60 (F2) [Wll] +AF58+AF59-AF51/3 AG60 (F2) [Wll) +AG58+AG59-AG51/3 AH60 (F2) [Wll] +AH58+AH59-AH51/3 AC62 [W5] ' MASS SUMMARY AD63 [W17] 'HA [kg] AE6 3 (F2) [Wll] tHF(AH7«"Yes"»AND»AH6»l,SG:$D$l J AD? 1 1K17] 'HSE [kg] J AE71 (F2) HU] + SG:$DS59,(AE86-AE6 j AK16 [K17] 'SFB [W/n2] j AL16 |WU] +SG:$D$61*(SG:$D$15/AL15)*2 j AM 16 [Wll] +$G:SD$61*(SG:SD$15/AM15)"2 j AN 16 [Wll] +SG:SD$61"(SG:SDS15/AN15)"2 j AO 16 [Kill +$G:$DS61*($G:$DS15/A015)*2 j AJ18 (W5) ' REFRIGERATION MASS PRELIMINARY CALCULATIONS j AK19 [K17J ' Oxidizer j AN 19 (Wll) ' Fuel j AK20 [K17) 'HOX [W/n2] j AL20 (F<0 (Wll) +$G:$DS50"SG:SDS21"SG:SDS6t*t j AN20 [Wll] 'HF j A020 (Ft) (Wll) +$G:SDS58a$G:$D$20*$G:$D$62*4 j AK21 1K17) 'ROXT |n/(kg)-l/3] j AL21 < F4) [Wll) (0.75/(SG:SD$18'SG:$DS50'(lPI))*(l/3) j AN 21 [Kill ' RFT j A021 (Ft) [Wll] (0.75/($G:$DS16"SG:SD$t9"@PI))"(l/3) j AK22 I W 17] •QIN [W/f kg)-2/3] j AL22 (Ft) [Will (AL16*(l-SG:SDS5t),SG:SDS2t)»gPI'AL21-2 j A X 22 1*11] 'QIN j A022 (FM [Wll] (A016'(l-SG:$DS3t),SG:SDS23)"(iPI*A021"2 j AK 2 3 [W17] 'Q0l ;T [ " " ] j AL23 (Ft) IWll) (AL20*SG:SD$32)"t"pPI«AL21-2 j AN 23 [Kill 'QOUT j A023 (Ft) [Wll) (A020"$G:SDS3l),t,@PI'A021"2 j AK2t IK17 1 'QCOOL I " "1 j AL24 (Ft) [Wll) +SG:SDS50*(AL22-AL23) j AN24 [Will 'QCOOL j A02t (Ft) [Will +$G;$DSt9*(A022-A023) j AK25 I W 1 7 J 'MRPRE [kg-1/3] j AL25 (Ft) (Wll) (SG:$DStt*AL2t/SG:SDS52) j AN 2 5 [Kill '. M RP R £ j A025 (Ft) (KUJ (SG:$DSt3*A02t/SG:$DS51) j A J 27 [W5| ' ROCKET DESIGN DETAILS (stage dependence) j AK28 [K17] 'FEED SYSTEM USED j AL28 [Ml| "pump j A?26 [Wll] " purap j AN28 [Kill "purap j A028 [Wll| "pump j AK29 [K17J •THRDES [kg] j AL29 (FO) [Will +SG:SDS66*AL86 j AM29 (FO) [Wll] +$G:SDS66*AM86 j AN 29 (FO) [Will + SG:SD$66*AN86 j A029 (FO) [Kill +SG:SDS66"A086 j AK30 (F2) [W17J •TIMEB [sec] j AL30 (F2) [Kill + A1 79 • A1 1 1 /AI 2Q j AM30 (F2) [Wll] gIF( AM29=0,0,AM79*AM13/AM29) j AN 30 (F2) [Wll] ^IF(AN29«0,0,AN79'AN13/AN29) j A030 (F2) [Wll] §IF(A029*0,0,A079"A013/A029) j AK31 [HI 7| 'OPT. FEED SYSTEH j AL31 [Wll] gIF(AL30<=(8130.1*AL29--0.75867)," press."," pump") j AM 31 [Kli) §IF(AM29«0," n/a",j!IF(AM3O<«(013O.l*AK29--0.75867)," pres s pump")) J AN31 : (Wll) piF(AN29=0," n/a",piF(AN30<=(8130.1*AN29--0.75867)," pres s pump")) J A031 ! [Wll] @IF(A029«0," n/a",0IF(A030<•(8130.l*A029*-0.75867),» pres s. pump")) J:AK32: (W17) 'PT (MPs) J: AL32: (F4 ) (Wll) £IF(AL28" 'PUMP50*0.101325/14.696,1. 5 *SC:SFS 28*0.101325/14. 6 96) J:AM32: (F4) (Wll) §IF(AM28= 'PUMP",50*0.101325/14.696,1.5"SC:SFS28•0.101325/14.6 96) J:AN32: (F4 ) (Wll) #IF(AN28» 'PUMP",30*0.101325/14.696,1.5"SC:$FS28*0.101325/14.6 96) J:A032: = "No" ,-(3*AM J: A06B: (F2 ) [Wll] piF(A078= 0 l0,piF(AL8»"Yes"ffAND#$C:SGS35»"No" J AK71 I w 1 7 J 'MSE [kg] J AL71 C F 2) *11] + SG:SD$59,(AL86-Al6 J 088 (Wll) +SCS1 J A91 [W5] 'MARS SURFACE DETAILS J A92 (W5) \* J B92 (Vi 17) \» J C92 [Wll] \- J D92 [Wll] \» J E92 [Wll] \« J F92 [Wll] \* J B93 [W17] 'MARS ROVER [kg] J C93 [Wll] +SG:SDSA2 J B9A [ W17] 'OX. FACTORY [kg] J C9A (F2) [Wll] +$E:SCS62 J D9A (Wll) ' > J E9A [Wll] 'MASS USED J F9 ^ (F2) |W11) 0IF(SXS8="Yes",C9A,0) J B95 [W17] 'FUEL FACTORY [kg] J C95 ( F 2) [Wll) 1008 J D95 [Wll] ' > J E95 (Wll) 'MASS USED J F95 (F2) (Wll) ()IF(SAA$8*"Yes"iC93iO) J B97 [ W 17) OXIDIZER REFRIGERATION UKITS [kg] J C98 (Wll] 'STAGE I J K98 |WU] "STAGE II J E98 (Wll) "STAGE III J F98 (tall) "STAGE IV J B99 (W17] "LEG 6 J C99 (FA) [Wll] +SF:S FS 116 J [>09 (FA) (Wll) + SF: SFS 11 5 J E99 (FA) (Wll) +$F:SFSUA J F99 (F ^) (Wll) +SF: SFS113 J B100 I W17 J "LEG 5 J C100 (FA) (Wll) +SF:SFS120 J D100 (FA) [Wll] +SF:SFS 119 J E100 (FA ) (till) +SF :SFS 11 8 J F100 (FA) [Wll] +SF: SFS 117 J B101 [W17) "LEG A J C101 (FA ) [Wll] +SF: SFS12A J 1)1 01 (FA) [Wll] +SF:SFS123 J E 1 01 (FA) [Wll] +SF:SFS122 .1 F 1 01 (FA) [Wll] +SFISFS121 J B103 [W17] "TOTAL J C10 3 (FA) [Wll] 0SUM(C99..F101) J D103 [Wll] ' > J E 103 [Wll] 'MASS USED J F 103 (FA) [Wll] £IF(SXS7="Yes",C103,0) J B105 [W17] 'FUEL REFRIGERATION UNITS [kg] J C 106 [Wll] "STAGE I J D106 [Wll] "STAGE II J E106 [Wll) "STAGE III J F106 [Wll] "STAGE IV J B107 [W17] "LEG 6 J C107 (FA) [Wll] +SF : SGS 116 J D107 (FA) [Wll] +SF:SGS 115 J E107 (FA) [Wll] +SF:SGS 11A J F107 (FA) [Wll] +SF:SGS 113 355 J B108 [ W171 -LEG 5 J C108 (FA) [Wll] +SF:SGS120 J D108 (FA) (Wll) +SF:SGS 119 J E108 (FA) [Wll) +SF:SGS118 J F108 (FA) (Wll) +SF:SGS117 J B109 [W17 ] -LEG A J C109 (FA) [Wll] +SF:SGS12A J D109 (FA) [Wll] +SF:SGS123 J E109 (FA) [Wll] +SF:SGS122 J F109 (FA) [Wll] +SF:SGS121 J Bill [W17] "TOTAL J CI 11 (FA) [Wll] gSUM(C107. J Dill [Wll] > > J Ell 1 [Wll] 'MASS USED J Fill (FA) [Wll] £IF(SAAS7» J B11 3 [ W I 7] 'OXIDIZER NECES J cm [Wll] "STAGE I J D114 [Wll] "STAGE II J E11A [Wll] "STAGE III J F11A [Will "STAGE IV J B11 5 I W 1 7] "LEG 6 J C11 5 (FA) [Wll] +C78 J D115 (FA) [Wll] +D78 J El 15 (FA) (Wll] +E78 J F115 J M106 CF1) [Wll] +SG:SDS26"M105 J 1107 [V17] 'MOX |kg] J J107 (FO) (WU) gSUM(X78..AA78)+@SUM(AE78..AH78)+gSUH(AL7B..A078) J L107 (Wll) •MOX [kg] J .1107 (F1 ) [Wll] gSUK(C78..F78)+jlSUM(J78. .M78)+§SUM(Q78. .T78)-F119 J HOG [W17] 'HOXL [kg] J jioe (Fl) [Wll] +SG:SDS27"(J107-AL78) J L108 [Wll] 'MOXL [kg] J M108 (Fl) [Wll] +SG:SD$28*M107 J 1109 [ W 17) 'MP [kg] J J109 (FO) [Wll] JISUM(X79..AA79)+gSUM(AE79. .AH79) +0SUM(AL79. .A079) J L109 (Wll) •MP [kg] J no9 (Fl) [Wll] gSUH(C79..F79)+£SUM(J79..M79)tgSUM(Q79..T79) J 1111 IW17) 'ML [kg] J Jill (FO) [Wll] gSUK(J92..J103) J LI 11 [Wll] 'ML [kg] J M.I 11 (Fl ) [Wll] gSUM(M92..M103) 359 J R88 [Wll] + SCS1 J 090 (W5 ] ' TOTAL MISSION J P91 1 W 1 7] •MA [kg] J Q91 (F2) (Wll) +J92+M92 J S91 [Wll] 'ME [kg] J T91 (F2) [Wll) +J98+M98 J P92 1 W17] 'MG [kg] J Q92 (F2) [Wll] +J93+M93 J S92 [Wll] 'MSE [kg] J T92 (F2) [Wll] +J99+M99 J P93 [ W17] 'MSS [kg] J Q93 (F2) [Wll] +J94+M94 J S93 [Wll] 'MFS [kg] J T93 (F2) [Wll] +J100+M100 j P94 [W1 7] 'MRF [kg] J Q9<< ( F 2 ) [Wll] +J95+M95+F111 J S94 1 W1 1 j •MS [kg] J 19". ( F 2 ) (Wll) +J101+M101 J P95 1 W 1 7] 'MROX [kg] J Q95 (F 2) [Wll) +J96*M96+F103 J S95 IWll) 'MFT [kg] J T95 (F2) [Wll] +J102+M102 J S96 [Wll] 'f.OXT | kg) J T96 ( F2 ) [Wll] •J103 + M103 .1 P97 1 W 17 ) •MF Ikg) J Q97 (F 2 ) [Wll] •J105+M105 J P98 1 W1 7 ) '. M.FL [kg] J Q98 (F 2 ) [Wll] +J106+M106 J P99 |H7| MOX [kg] J Q99 IF;) [Wll] +J107+M107 J PI 00 |»17] 'MOXL [kg) J 0100 (F2) [Wll] +J108+M108 J SI 00 [Wll] 'MSAKPL [kg] J T1 00 (F 2) [Wll] + SG:SDS<»5 J P101 [W17 1 '.MP [kg] J Q101 I F 2 ) [Wll] tJ109+M109 J S101 [Wll) 'MSUPP [kg) J T10 1 (F 2) [Wll] tSG:SDSA6 J P102 1 W17) 'ML [kg] J Q102 (F 2) [Wll] gSUM(Q91..T96) J S 102 [Wll) 'UROV [kg) J T102 (F2) (Wll) + SG:SDS42 J S 103 [Wll) 'MOXF AC [kg] J T10 3 ( F 2) (Wll) +F94 J SIC. [Wll) 'MFFAC [kg] J TIC. (F2) (Wll) +F95 J S106 [Wll] \- J T106 IW11] \- J S107 (Wll) 'MLEO (kg) J T107 (F2) [Wll] gSUM( Q91.. Q97 ) +Q99+£SUM (T91. .T10<() A1 (W5] 'SHEET »U CI (Wll) +SC:SASA2 D1 [Wll] +C:Ct2 D2 [Vll] 'MLEO [kg] E2 (FO) [Wll] +T107 A't [W5] +SB:SBSA7 Ci (Wll) +SB:SCS'i7 A5 (W5] \- B5 IW17] C 5 [Wll] D5 [Wll] \» E5 [Wll] \= F5 [Wll] \« B6 [W17] 'TOTAL DELTA V C 6 [Wll] + SB:SGS B16 7 'SFB (W/n2] C16 1 +SG:SE$61"(SG:SES15/C15)'2 D16 I +SG:$ES61,(SG:$E$15/D15)"2 E16 1 +SG:SES61"(SG:SES15/E15)"2 F16 1 +SG:SES61*(SG:$ES15/F15)"2 A18 ] ' REFRIGERATION MASS PRELIMINARY CALCULATIONS B]9 7 ' Oxidizer E19 1 ' Fuel B20 7 'HOX [W/n2] C20 ) (Wll) +$G:SES58"$G:$ES21*SG:SE$65"4 E20 1 •HF F20 ) [Wll] +$G:SES58"SG:$E$20*$G:$E$63*'i B21 7 'ROXT lm/(kg)"1/3] C 21 ) (Wll) (0.75/($G:$ES19"SG:$ES30"@PI))*(l/3) E21 1 ' RFT F21 ) [WU] (0.75/(SG:SES17"SG:SES'i9,PPI))"(1/3) B22 7 •QIN [W/(kg)"2/3] C 2 2 ) [Wll] (C16,(l-SG:$ES54)*$G:SES2 F(D32>=SD:SFS7114,piF(D32>»SD:SES71,3,SD:SJ$61))))) K:E33: (FO) (Wll) piF(E32> = SD:SIS71,7,piF(E32>»SD:SHS71,6,piF(E32>=SD:SGS71,5.pI F(E32>=SD:SFS71,41piF(E32>«SD:SES71,3,SD:SJS61))))) K:F33: (FO) (WU) pIF(F32>=SD:SIS71,7,pIF(F32>=SD:SHS71,6,pIF(F32>«SD:SGS71,5,pI F(F32>=SD:SFS71,4,piF(F32>«SD:SES71,3,SD:SJS61))))) K:B34: IV. 1- 'VOX (m3/kg] K:C 34: (Wll) piNDEX(SD:SBS71 SD:SIS87,C33,SD:SFS125) K:D34 : I v. 1 1 J PINDEX(SD:SBS71 SD:SISB7.D33,SD:SFS125) K : E'J'I : (Wll] piNDEX(SD:SBS71 SD:SIS87,E33.SD:SFS125) K:F3<-: (Wll) piS'DEX ( S D: SBS7 1 SD:SIS87,F33,SD:SFS125) K:B35: (W 17 'VF (m3/k g] K:C35: 1*11] PINDEX(SD:SBS71 SD:SIS87,C33,SD:SFS124) K:D35: (Wll] pINDEX(SD:SBS71 SD:S1S87,D33,SD:SFS12'«) K:E35: (Will PINDEX(SD:SBS71 SD:SIS87,E33,SD:SFS124) K:F 35 : (Wll] PINDEX(SD:SBS71 SD:SIS87,F33,SD:SFS124) K:B36: |W17] ' OXTC K:C36: F 4) (Wll] piF(C28="pres s , 3* SG :SES82*SG:SE$8«.*C32* 1000000*C34"(l-C14)/( 2•SG:SES65"(l + (1/SC:SFS56))),3•SG:SES83•SG:SES84"C32•1000000*C34*(1-C14)/(2"SG: S ES85*(l + (1/SC:SFS56) ) ) ) K :1)36 : ( F4 ) (Wll] pI F (D28 = "pres s . " , 3" SG :SES82 • SG: S ES84 "D32 * 1000000-D34 • (1-D14 ) /( 2*SG:SES85*(1+ (1/SC:$FS56))),3*SG:SES83•SG:SES84•D32•1000000•D34•(1-D14)/(2*SO:S ES85"(l + (1/SC:SFS56) ) )) K :E36 : (F<<) (Wll) §IF (E28 = "pres s . " ,3 • SG :SES 82 • S G: S ES 84 «E32 • 1000000 'E34 * (1-E14 ) /( 2,SG:SES85,(1+ (1/SC:SFS56))),3"SG:SES83•SG:SES84«E32•1000000•E34"(1-E14 )/(2"SG: S E S 6 5 *(l+ (1/SC:SFS56) ))) K : F 36: (F4) [Wll] piF(F28= "press3"SG:SES82•SG:SES84*F32* 1000000 "F34•( 1-F14)/( 2*SG:SFS85*(1+ (1/SC:SFS56))),3*SG:SES83•SG:SES84"F32•1000000 *F34•(1-F14)/(2•SG: S ES85-<1*(1/$C:SFS56) ) )) K : B37 : [W17] 'FTC K:C37 : (F4) [WU] piF (C28= press. ,3*SG:SES76"SG:SES78*C32*1000000*C35,(l-C14)/( 2,SG:SES79*(1+SC:SFS56)),3'SG:SES77*SG:SES78"C32*1000C00*C35"(1-C14)/(2•SG:SES79 "(1+SC:SFS56))) K:D37: (F4) [Wll] piF(D28«"press3*SG:SES76*SG:SES78*D32*1000000*035*(1-D14)/( 2*SG:SES79*(1+ SC:SFS56)),3"SG:SES77"SG:SES78*D32* 1000000*D35*(1-D14)/(2"SG:SES79 • (1+ SC:SFS56))) K:E37: (F4) [Wll] piF(E28="press3*SG:SES76*SG:SES78*E32* 1000000*E35*(1-E14)/( 2,SG:SES79"(1+SC:SFS56)),3,SG:SES77"SG:SES78*E32*1000000*E35*(1_E14)/(2*SG:SES79 *(1+ SC: SFS56) ) ) K:F37: (FA) [Wll] piF(F28="press.",3*SG:SES76*SG:SES78*F32*1000000*F35'(1-F14)/( 2*SG:SES79,(1+$C:SFS56)),3*SG:SES77*SG:SES78"F32*1000000*F35*(1-F14)/(2"SG:SES79 • (1+ SC: SFS56) ) ) K:B38: [W17] 'FSC K.C38: (FA) (Wll] §IF(C28«"pressSG:$ES29«SC:SFS57/500,SG:SES30»(1EXP(0.000759' SC:SFS57)) K:D38: (FA) [Wll] piF(D28="pressSG:SE$29*SC:SFS57/500,SG:SE$30"#EXP(0.000759' SC:SFS57)) K :E38 : (FA) [Wll] 0IF(E28«"pressSG:SES29"SC:SFS57/500,SG:SES30'§EXP(0.000759" SC:SFS57)) K :F38 : (FA) [Wll] @IF(F28*"press.",$G:SES29'SC:SF$57/500,$G:SE$30*gEXP(0.000759" SC:S FS57)) K AAO: [W 5) ' INTERMEDIATE MASS CALCS K BA1: [W17] 'COEFSUM K CA 1 : (FA) [Wll] +SG:SES35+SG:SESA8+C37+C36+SG:SES59+SG:SES60+C38+SG SES22 K DAI : (FA) [Wll] +SG:SES35+SG:SESA8+D37+D36+SG:SES59+SG:SES60+D38+SG SES22 K EA1 : (FA) [W11j +SG:SES35+SG:SESA8+E37+E36+SG:SES59+SG:SES60+E38+SG SES22 K FAl: (FA) [Wll] +SG:SES35+SG:SESA8+F37+F36+SG:SES59+SG:SES60+F38+SG SES22 K BA2: [ W1 7] 'COEFTL K C A 2: (FA) [Wll] 0IF(F7 = "YES"«AND//F6=1 , SG: SES1A + CA1 , CA1) K DA 2: (FA) [Wll) 0IF(F7="YES"»AND«F6*2,SG:SES1A + DA1,DA 1) K EA 2: (FA) [Wll] 0IF(F7="YES"«AND»F6=3,SG:SES1A+EA1,EA1) K F A 2: (FA) [Wll] £IF(F7n" YES "//AND(/F6 = A , SG :SES 1A + FA 1 , FA 1) K BA 3 : [ W17 1 'C0EFTL1 K CA3 : (FA) [Wll ] PIF(C7 = "N0" ,0,#IF(F7»"YES"»AND»F6 = 1,SG:SES1A + SG:SES35+SG :SES60 .SG:SES 35+SG:SES60)) K :DA 3 : ( FA ) [Wll] 01F(C7 = "NO",0,0IF(F7="YES"wANDwF6*2,SG:SES1A + SG:SES35+ SG:SES60 , SG : SES35-SG:SES60)) K : EA 3 : (FA) [mi] piF(C7 = "NO",0.§1F(F7="YES"»AND«F6«3,SG:SES1A + SG:SES35+ SG:SES60 .SG; SES 35 +SG:SES60)) K :F A 3 : (FA) [Wll] 01F(C7 =" NO" , 0 , £IF( F7= " YES " »AND/F B60 IW17] •MF [kg] C60 (FA) (Wll) +C58+C59-C51/3 D60 (FA) (Wll] +D58+D59-D51/3 E60 (FA) (Wll] +E58+E59-E51/3 F60 (FA) (Wll) +F58+F59-F51/3 A62 [ W5] ' MASS SUMMARY C62 (Wll) '(Hohmann to highly eccentric Earth orbit) B63 [W17] 'MA [kg] C63 (F2) {Wll eiF(F7-"Yes"//AND#F6«l , SG:SES1A • (C86-C6A ),0) D6 3 (F2) [Wll @IF(F7»"Yes "ff AND//F6«2.$G:SES1A*( D86-D6A ),0) E63 (F2) (Wll tUF(F7»"Yes"0AND#F6«3,$G:SESlA*(E86-E6A),0) F63 (F2) [Wll 0IF(F7- Yes"i/AND#F6«A,SG:$E$1A*(F86-F6A),0) B6A (W17] •ML [kg) C6A (F2) [Wll + C11 D6A (F2) [Wll + D11 E6^ (F2) [Wll + E11 F6A (F2) [Wll + F11 B65 [W17] 'MG [kg] C65 (F2) [Wll +SG:SES35«(C86-C6A)+0,A D65 C F 2) (Wll +SG:SE$35*(D86-D6A)+ 0. A E65 (F2) [Wll +SG:SES35"(E86-E6A)+0.A F65 (F2) [Wll +SG:SES35'(F86-F6A)+0.A B66 (W17) •r.ss [kg] C66 (F2) [Wll + SG.'SES601 (C86-C6A) D66 !F 2) [Wll +SG:SES60*(D86-D6A) E66 (F2) (Wll +SG:SES60'(E06-E6A) F66 f F 2) (Wll +SG:SES60"(F86-F6A) B67 [WIT] ' MRF [kg] C6" ( F 2) (Wll @1F(F8="Yes ",F25'C77*(2/3),0) D67 ( F 2 ) [Wll piF(F8""Yes"iG25'D77"(2/3)>0) E6: (F 2) (Wll @IF(F8="Yes",H25"E77"(2/3).0) F6" ( F2) [Wll PI F(F 8•" Y e s " iI 25 • F 77 '(2/3) t0) BfcS [ W 1 7] '"ROX [kg] C68 ( F 2) [Wll £IF(C8 Yes",C25"C78'(2/3),0) D68 I F 2 ) [Wll £IF(C8="Yes",D25*D78*(2/3).0) E68 (F2) [Wll £IF(C8="Yes",E25"E78*(2/3),0) F66 (F2) [Wll piF(C8="Yes". F2VF78"(2/3),0) B70 [M?| ' ME i kg] C70 (F2 ) [Wll +SG:SES22'(C86-C6A) D70 (F2 ) (Wll +SG:SES22'(D86-D6A) E70 (F 2) [Wll +SG:SES22'(E86-E6A) F70 ( F 2) [Wll +SG:SES22"(F86-F6A) B7 1 [W17] ' MSE [kg] C 7 1 ( F2 ) [Wll +SG:SES59"(C86-C6A) D7] ( F2) [Wll + SG:SES59*.(D86-D6A) E71 ( F2) [Wll +SG:SES39"(E86-E6A) F7 1 (F2) (Wll +SG:SES59'(F86-F6A) B72 [W17] ' MFS [kg] C72 (F2) (Wll +C38•(C86-C6A) D72 (F2) [Wll + D38 *(D86-D6A) E72 (F2) [Wll + E38*(E86-E6A) F72 (F2) [Wll + F38"(F86-F6A) B73 (V.17] •MN [kg] C73 ( F2) (Wll +SG:SESA7'(C86-C6A) D7 3 (F2) [Wll +SG:SESA7*(D86-D6A) 366 K E73 F2) [Wll] +SG:SES47«(E86-E64) K F73 F2) [Wll] +SG:SES ,, K:E86: (F2) (Wll) $1F(E14b1,0I9IF(C7> NO"IE79/(1-E14)IE64+E82/(1-E43))) K:F06 : 116 (W17 'SFB [W/n2] J16 (W11 +SG:SES61"(SG:SES15/J15)~2 K16 [Wll +SG:SES61,(SG:SES15/K15)"2 L16 [ W11 +SG:SES61*(SG:SES15/L15)"2 M16 Ikll +SG:SES61"(SG!SES15/M15)*2 H18 |W5] ' REFRIGERATION MASS PRELIMINARY CALCULATIONS 119 [W17 ' Oxidizer L19 [ W11 • Fuel 120 [ W1 7 'HOX [W/m2] J20 (FA) [Wll] +SG:$E$5B*SG:$ES21"SG:SES65*A L20 [Wll 'HF M20 (F^) [Wll] +SG:SES58,SG:SES20'SG:SES63*A 121 [ W17 'ROXT (•/(kg)*1/3] J21 (FA) [Wll] (0.75/(SG:SES19*SG:SES50"gPI))"(1/3) L21 [Wll ' RFT u. 21 (FA) [Wll] (0.75/(SG:SES17-SG:SESA9'gPI))"(1/3) 122 [WIT 'QIN [W/(kg)"2/3] J22 (FA) [Wll] (J16"(l-SG:SESJA)"SG:SES2A)ippi"J21*2 L22 [Wll 'QIK "22 (FA ) [Wll] (Ml 6"(1-SG:SES3A)'SG:SES23)"gPI*M21"2 123 [ W1 7 1 QOL'T [ " " ] J 23 (FA) (Wll] (J20,SG:SES32)"A,pPIiJ21"2 L2 3 [Wll ' QOL'T M23 (F-) IVf 11J (M20"SG:SES31 )*A"0PI'M21"2 I 2h [ W 1 7 •QCOOL I " " ] J2A (Fa) [Wll] +SG:SES50*(J22-J23) L2& (Wll 'QCOOL HTJj (FA) (WllJ. +SG:SESA9'(M22-M23) 125 [ W 1 7 ' MRPRE [kg*1/3] J25 (FA ) [Wll] (SG:SESAA•J2A/SG:SES52) L25 I W 1 1 ' MRPRE w 2 3 (FA) [Wll] (SG:SESA3"M2A/SG:$ES51) H27 [<•5) ROCKET DESIGN DETAILS (stage dependence) I 2B [WIT 'FEED SYSTEM USED J 2 8 [Wll "press. K28 [Wll "press. L28 (Wll "press . ".2 8 (Wll "press. 129 (W 17 'THRDES [kg] J29 (F0) [Wll] + SG:SES 66"JB6 K29 (FO) [WllJ +SG:SES66*KB6 L29 (FO) (Wll] +SG:SES66 *L86 ^29 (FO) (Wll] +SG:SE$66"MB6 I 30 (F2) [W17] 'TIMEB (sec) J30 (F2) [Wll] +J79*J13/J29 K 30 (F2) [Wll] £IF(K29=0,0,K79*K13/K29) L30 (F2) (Wll) gIF(L29»0,0,L79'L13/L29) M 30 (F2) [Wll] 0IF(M29«O,O,M79'M13/M29) 131 [ W17] •OPT. FEED SYSTEM J31 [Wll] £lF(J30<*(8130.1"J29"-0.75867)," press punp") K 31 [Wll J @IF(K29=0," n/a ^IF(K30<«(8130.l'R29--0.75867), " press, pump" )) K :L 3 1 : [Wll] @IF(L29=0," n/a",#IF(L30<-(0130.l»L29"-0.75867)," press. pump" )1 K :«31! [Wll] pi F(.".29 =0, " n/a",£IF(M3O<»(013O.1'M29*-O.75867)," press. pump" )) K:: 132: (W17] 'PT (MPa) K::J32: (F K 160 [ W17] 'MF [kg] K J60 (F<0 Wll] +J58+J59-J51/3 K K60 (F4 ) Wll] +K58+K59-K51/3 K L60 (F«) Wll] +L58+L59-L51/3 K M60 C F K P16 [ W17 ] 'SFB [W/M2] K Q16 [Wll] +$G:$ES61,(SG:SE$15/Q15)"2 K R16 [Wll] +SG:SES61*(SG:SES15/R1S)*2 K S16 (Wll) +SG:$E$61"(SC:$ES15/S1S)"2 K Tib [Wll] +SG:SES61"(SG:SES15/T15)*2 K 018 [W5] ' REFRIGERATION HASS PRELIMINARY CALCULATIONS K P19 [ W17] 1 Oxidizer K S19 (Wll) ' Fuel K P20 (W17] 'HOX [W/n2] K Q20 (F<0 (Villi] +SG:SES58"$G:SES21*SG:SES65-^ K S 20 [Wll] 'HF K T20 (F K P60 [W17] 'HF [kg] K Q60 (F1) [Will +Q58+Q59-Q51/3 K R60 (F4) [Wll] +R58+R59-R51/3 K S60 (F4) IK11) +S58+S59-S51/3 K T60 (FM [Wll] +T58+T59-T51/3 K 062 [W5] ' MASS SUMMARY K P63 [ W17 ] 'MA [kg] K Q63 (F2) [Wll] 0IF(T7«"Yes"»AND«T6= l,SG:SES14»(Q86-Q64) ,0) K R63 (F2) [Will @IF(T7= "Yes"«AND»T6"2.SG:SESl'i"(R86-R64),0) K S63 (F2) [Wll] £IF(T7="Yes"»MND«T6«3,SG:SES14 *(S06-S64),0) K T63 (F2) [Wll] {UF(T7»"Yes"#ANDffT6»4,SG:$ES14"(T86-T64),0) K P64 [ W17) •ML (kg) K Q6A (F2) (Wll) +Q11 K »6M IF2) [Wll] •R11 K S64 ( F2 ) [Wll] +S11 K T6<< (F2) [Wll] +T11 K P65 Ifcl? J ' MG [kg] K Q65 ( F2 ) IW11] +SG:SES35'(Q86-Q64)+0.4 K R65 (F2) (Will +SG : SES 35 • (R86-R64)+0 . 4 K S65 ( F 2 ) [Wll] +SG:SES35*(S86-S64)+0.4 K Tb5 ( F 2 ) t f1 1 ] •SG:SES35'(T86-T64)+0.4 K P66 1-17) • MSS [kg] K Q66 ( F 2 ) [Wll] +SG:SES60*(Q86-Q64) K Rb6 ( F 2) [Wll] •SG:SES60"(R86-R64) K S 66 ( F 2 ) [Will +SG:$ES60"(S86-S64) K T66 ( F 2) (Wll) +SG: SES60"(T06-T64) K P6T |W17| • ?.RF 1 kg] K Q67 ( F 2 ) (Will piF(T8="Yes".T25-Q77-(2/3).0) K R67 < F 2) [Wll] £IF(T0""Yes",I 25 *R77 "(2/3),0) K S67 1 F 2) [Will @IF(T0="Yes",V25*S77"(2/3),O) h Tb" ( F 2 ) [Will piF(T8«"Yes",W25"T77"(2/3),0) K P68 (V.17) •^ROX [kg] K Q68 ( F2) [Wll] PIF(Q8«"Yes",Q25"Q70-(2/3),0) K R68 ( F 2 ) [Wll] piF(Q8="Yes",R25*R78"(2/3),0) K S6e i. F 2) iwu] piF(Q8="Yes".S25 *S7B"(2/3),0) K TiS I F2) [Wll] 0IF(Q0="Yes",T25*T73*(2/3),0) K F70 11.17] 'ME [kg] K Q70 IF 2) [Wll] pIF (Q7=" No" , SG: SES22* (Q06-Q64) ,Q81 * SH: SES7 ) K R70 lF2) [Wll] £IF(Q7«»"No"tSG:SES22"(R06-R64),R01*SH:SES7) K S70 (F2) [Wll] 0IF(Q7= "No" , SG : SES22" (S06-S64) .S81"SH: SES7) K T70 ( F 2) [Wll] £IF(Q7 = "No"iSG:SES22*(T86-T64),T81,SH:$ES7) K P71 IW17J •USE [kg] K Q71 (F2) [Wll] PIT(Q7="No",SG:SES39'(Q86-Q64),Q81*SH:SES8) K R71 ( F2) [Wll] @ IF(Q7«"N o " ,SG:SES59*(R06-R64),R01"$H:SES0) K S71 (F2) [Wll] eiF(Q7= "No",SG:SES59*(S06-S6 , 4 K T73 F2) [Wll] No"tSG:SFS K H15 : [¥17 'RBS [AU] K *15: [Wll +SB:SCS6 K V 15 ! [¥11 +SB:SCS6 K Z15: [Wll +SB:SCS6 K AA15 [ 1 ) +SB:SCS6 K ¥16: [ W17 'SFB (W/m2] K X16 : [Wll •SG:SES61"(SG:SES15/X15)'2 K V16 : [Wll +SG:SES61*(SG:SES15/Y15)*2 K Z 16: (¥11 +SG:SES61*(SG:SES15/Z15)"2 K AA16 [¥1 J +SG:SES61"(SG:SES15/AA15)* 2 K V 18 : [W5J ' REFRIGERATION HASS PRELIMINARY CALCULATIONS K ¥19: [¥17 Oxidizer K Z 19: [Wll ' Fuel K ¥20 : f ¥r 17 ' HOX [¥/m2] K X20: (F<.) [Wll] + SG:SES50iSG:$ES21,SG:$ES6 K AA30 (F2) [Wll| £IF(AA29=OtOiAA79*AA13/AA29) K ¥31 : 1 ¥ 17 | 'OPT. FEED SYSTEM K X 31: 1 ¥ 1 1 ] @IF (X30<«( 813.0. l*X29"-0. 75067 ) , " press. " , " K: V31: [Wll] 0IF(Y29-0, " n/a",0IF( Y30<»(8130.1»>'29"-0.75867)," press.", pump"))niimn" 1 1 K :Z31: [Wll] 0IF(Z29=O," n/a",0IF(Z30<=(8130.1*Z29*-0.75867),» press.", pump"))mimn'Mt K:AA31: [Wll] 0IF(AA29*O, n/a",eiF(AA30<«(8130.l*AA29--0.75867)," pres s."," pump")) K :W 3 2 : (W17] 'PT [HPa] K :X32 : (F4) [Wll] 0IF(X28»"PUMP" ,50*0 .101325/l^t. 696,1. 5* SC: SFS48*0. 101325/14 .696 ) K : Y32: (F4) (Wll) 01F(Y28»"PUMP",50*0.101325/14.696,1.5"SC:SFS48-0.101325/14.696 ) K :Z 32 : (F4) [Wll] 0IF(Z28= "PUMP",50*0.101325/14.696,1.5 *SC:SFS48-0.101325/14.696 ) K:A A 3 2: (F4) [Wll] 0IF(AA28="PUMP",50*0.101325/14.696,1.5"SC:SFS48'0 .101325/14.6 96) K:W33: [K17] 'PINDEX K :X 3 3: (F0) [Wll] #11F ( X32 > = SD: S IS71 ,7 ,0IF( X32> = SD: SHS71 ,6 ,0IF (X32> = SD: SGS71 ,5 .pi F(X32>=SD:$FS71,4,0IF AD16 |W17] 'SFB [W/m2 J AE16 (till I +SG:SES61'(SG:SES15/AE15)*2 AF16 |W11| +SG:SES61*(SG:SES15/AF15)"2 AG 16 [Wll] +SG:SES61*(SG:$ES15/AGL5)*2 AH16 |W11] +SG:SES61"(SG:SES15/AH15)*2 AC 1 0 [M5] ' REFRIGERATION MASS PRELIMINARY CALCULATIONS AD19 [W17] • Oxidizer AG 19 [Wll] • Fuel AD20 [W17] 'HOX (W/m2] AE20 (FA) [Kill +SG!SES58'$G:SES21'SG:SES6«-« AG20 [Wll] 'HF AH20 (FA) [Wll] +SG:SES50*SG:SES2O,SG:$ES62"'t AD21 IW17] ' ROXT [m/(kg)"1/3) AE21 (FA) [Wll] (0.75/(SG:5ESia*SG:SES30*ePI))'(1/3) AG 21 [Wll] 'RFT AH21 ( F^ } [Wll] (0.75/($G:$E$16*$G:$E$A9"@PI)) (1/3) AD22 I W 17 ] ' QIS" [K/(kg)-2/3J AE22 (FA J [Wll] (AE16*(1-SG:SES5A) $G:$ES2A),@PI*AE21"2 AG22 [Wll] ' QIN AH22 (F<< ) I K I 1 ] (AH16"(1-SG:SES3A)' SG:SES23)"PPI,AH21*2 AD23 IK 17 J 'QOl'T 1 AE23 ( FA ) [K11J (AE20'SG:$ES32),A,@PI"AE21"2 AG 2 3 [Wll) 'QOL'T AH 2 3 (FA ) (Wll) (AH20"$G:SES31)"A,@PI"AH21"2 AD2A IW17| QCOOL I " " ) AE2A (FA) (Wll) +SG:SES50*(AE22-AE23) AG2A [Wll) 'QCOOL AH2A (FA) (Wll) +SG:SESA9*(AH22-AH23) AD25 [WIT] '."RPRE [kg" 1 / 3] AE25 (FA) j W11] (SG:SESAA"AE2A/SG:SES52) AG 25 |W11] 'KRPRE AH 2 5 (FA) [Wll] (SG:SESA3-AH2A/SG:SES51) AC 27 IK5 ] ROCKET DESIGN DETAILS (stage dependence) At)2 0 i» 1 ~ ] FEED SYSTEM ISED AE2H |WU] pump M 28 [Wll] pump AG 2 a I W 1 1 ] pump A H 28 (Wll] pump AD29 I w 1 7) THRDES (kg] AE 2^ (FO) (Wll tsG:SES66"AE86 AF29 (FO) (Wll +SG:SE$66*AF86 AG 29 (FO) (Wll +SG:SE$66'AG86 AH 24 (FO) (Wll +SG:SES66"AH86 AD3U (F2) [K17 'TIHEB [sec] AE 30 (F 2) [Wll +AE79*AE13/AE29 AF30 (F2) [Wll piF(AF29 = 0.0t AF79 *AF13/AF29) AG 30 (F2) [Wll PIF(AG29=0,0,AG79«AG13/AG29) AH 30 (F2) (Wll PIF(AH29»0>0,AH79"AH13/AH29) AD 31 IW17] 'OPT. FEED SYSTEM AE 3 1 (Wll) 0IF AE30< =(8130.1*AE29*-0.75867) , " press. pump") AF31 |K11) §IF AF29 =0," n/a",0IF(AF3O<-(8130.l"AF29*-0.75867)," pres pump") AG 3 1 : [Wll] 0IF AG29 =0, " n/a",£IF(AG30<=(6130.1«AG29*-0.75867)," pres pump") AH31 : (Wll) 0IF AH29-0," n/a",$IF(AH30<«(8130.l*AH29"-0.73867)," pres pump") K : A D 32 : (Vi 17 ] 'PT (MPa] K: AE32: (F4) [Wll] £IF(AE28*"PUMP".50*0.101325/14.696,1.5*SC:SFS48"0.101325/14.6 96) K:AF32: (FA) [Wll] eiF(AF28»"PUMP",50'0.101325/14.696,1.5»SC:SFS48«0.101325/14.6 96) K : A G 32 : (F4) [Wll] £IF(AG28«"PUMP",50"0.101325/14.696,1.5*SC:SF$48"0.101325/14.6 96) K: AH32: (F4) [Wll] £IF(AH28«"PUMP",50*0.101325/14.696,1.5"$C:$FS40*O.101325/14.6 96) K:AD33: (W17] 'PINDEX K:AE33: (FO) [Wll] @IF(AE32>*SD:SI$71,7,@IF(AE32>*SD:$H$71,6,@IF(AE32>'$D:SGS71, 5,@IF(AE32>n$D:SFS71,4,$IF(AE32>«SD:SES71,3,SD:SJS61))))) K:AF33: (FO) [Wll] $IF K AD71 [ W17 ] 'USE [kg] K AE71 CF2) (Wll] +SG:SE$59"(AE86-AE6M K AF71 IF2) IVill] +SG!SES59*(AF86-AF64) K AG71 (F2) [Wll] +$G:$E$59*(AG86-AG64) K AH71 (F2) [Hill +$G:SES59'(AB86-AH6<0 K AD7 2 [W17] 'MFS [kg] K AE72 (F2) [Wll] +AE30*(AE86-AE6A) K AF72 (F2) [Wtl] + AF38*(AF86-AF6'i) K AG7 2 (F2) [Mil] +AG38*(AG86-AG6<0 K AH72 (F2) 1X11) + AH38*(AH86-AH6 K AK16 'SFB IW/m2] K AL16 +SG:SES61»(SG:SES15/AL15)"2 K AM16 +SG:SES61"(SG:SES15/AM15)*2 K AN 16 +SG:SE$61*(SG:$ES15/AN15)*2 K AO 16 +SG:SES61"(SG:SES15/A015)*2 K A J1 8 • REFRIGERATION MASS PRELIMINARY CALCULATIONS K AK19 • Oxidizer K AS 19 ' Fue 1 K AK20 'HOX [W/d2] K AL20 [Wll] +SG:SES58"SG:SE$21,SG:SES6A"A K AS 20 •HF K A020 [Wll] •SG:SES56"SG:SES20"SG:SES62*'i K AK21 'ROXT (n/(kg)"1/3] K A L2 1 (Wll) (0.75/(SG:SES18"$G:SES30"(IPI))"(l/3) K *S21 RFT K *021 (Will (0.75/(SG:SES16"$G:SE$49,{iPI))"(l/3) K AK22 • QIN lW/(kg)"2/3] K AL22 [Wll] (AL16,(l-SG:SE$5<0"SG:SES2 K: AK38: (til?) 'FSC K: AL38 : (F4) [Wll) @IF ( AL28 = "press SG:SES29 • SC : SFS46/500, SG: SES30 "#EXP(0. 00075 9*SC:SFS48 )) K : AM38: (F4) [Wll] $IF(AM28»"press SG:SES29* SC: SFS48/500 , SG: SES30 '#EXP( 0. 00075 9"SC:SFS48)) K:AN38: (F4) [Wll] £IF K AK71 IW17} 'USE [kg) K A L71 (F2) [Wll| + $G:SE$59"(AL86-AL6 K F115 (FA) (Vill) + F7 8 K B116 [ W 17| "LEG 5 K C116 (FA) (Will + J78 K D116 (FA) (Wll) +K78 K EL 16 (FA) IWll] +L78 K F11 6 (FA) (Will +M78 K B117 C W 17 ] -LEG A K CI 17 (FA) (Wll] +Q78 K D117 (FA) [Wll] +R78 K E117 (FA) IWll] + S78 K F117 (FA) [Wll] +T78 K B1 19 [V17] •TOTAL K C119 (FA) (Wll) PSUM(C115. K DU9 1*11] • > K EL 1 9 [Ml] •MASS USED K F1 19 (FA) IWLL] 0IF(SXSB«" K B121 (*17) 'FLEL NECESSARY K C1 22 IWll] "STAGE I K D122 IWll! "STAGE 11 K El 22 IWll I "STAGE 111 K F 1 22 (Will "STAGE IV K B1 23 1 W 17 ] •LEG 6 K R: 23 (FA) IWll] + C77 K D! 23 (FA) |W11 j + D77 K EL 25 (F 4 ) IWll] •E77 K F 1 2 3 (FA) [Wll] + F77 K B1 24 1 W 1" J "LEG 5 K C 1 24 (F<< J (Will +.377 K D1 2" (FA) [Wll] + K77 K EL2<< (FA) IWLL] + L77 K F 1 2A (FA) IWll] + *77 K B1 2 f. 1 W 1 7 ] • LEG A K C 1 2D (F-) [Will + Q77 K D1 25 < F A ) [ W 1 1 i 1-R77 K El 25 (Ft) [Wll] + S77 K F 1 2 5 ( F4 ) IWLL) + T77 K B1 27 1*17] "TOTAL K C 1 27 t F« ) IWLL) PSUH(C123. .F125) K D127 IWll] ' > K El 27 (Will ' F ASS USED K F1 27 (FA) IWll] £IF(SAAS8= •Yes".C127.0) K C 1 20 (FA) 1W11J "STAGE I K U!28 (FA ) IWll] "STAGE II fi E I J8 (FA) IWll] "STAGE III K F 1 2b (FT) IWll] -STAGE IV K B129 (W 1 7] "LEG 6 K C 129 (FA) [Wll] +C03 K D129 (FA) IWll] +D83 K E1 29 (FA) [Wll] +E83 K F 12 9 (FA) [Wll] + F03 K B1 30 IW 17 J •LEG 5 K C 1 30 (FA) IWll) + J83 K D1 30 (FA) IWll] +K83 K El 30 (FA) |W11) + L83 K F 1 30 (FA) IWll] •M83 406 K:B131: (W17| * LEG 4 K: C131 : (F«) [k'llj +Q83 K: D131 : (F<.) |WU] +R63 K:El 31: (F4) [Wllj +S83 K: F131: ( F — . «_ »* - *• • r • r • r - '-** + -H - *: • *• - «! - r • r - ** - *• - :c :« h- ^ ;c »- .< . :e—*»-3»-3^3»-3^3 — 3 o rc»—*:c •-> it »—• rr 3 — rr •— :c »- X •— :« CO c/> m UOm •— *• *H 0 O T1 H- *ri t-' 3S — *- -n *- "0 — m — m-— — — O w "5 «— -rj — TJ <—— Kjl — Co W W » so v» )• X — X — -H — •—1 C/l — CO 1™ —» X „ -. «OMMP -H H 1S>_____ — TS ^ 7^ 'B) Tb._ 16—- 'a — 16<—• *B>.— «B> TT'B)?T15> 7^16yr c/i I , + ,—.'BJTRFO r- 1— t— I— ) ) ) r- t- r-) - • • -) r- r r- r— -4 -4 6 f- 05 H -* 0 07 -H 0 •H o«3 -sj 7 -si •sj CD 6 -si 6 •H H -1 •H -4 -4 O O O 6 -4 CN 0 O u* 0 CO6 -sj •sj -sj 0 >0 • • • - -J -4 co ~4 CD 0 0 0 O -4 O O Ui CO -J • CJ» • CO A K> • s_» . - s_* . s_<- . s_^ . A A A A . > V > 1 > » 3" 0 O O 0 0 •s| «4 o> ) O H- -sj -4 -4 «sj 3 O 3 -J w X- KJ •u o•sj 408 K M106: (F1 ) [Wll] +SG:SES26'M105 K 1107: 1*17) 'MOX [kg] K J107: (F0) (Wll) gSl'M(X78..AA78)+£SUM(AE70..AH78)+@SUM(AL70.. AO70) K L107: [kll] •MOX [kg] K n 107: C F1 ) [Wll] @SUM(C78..F78)+£SUM(J70. . M70)+@SUM(Q70..T70) -F119 K 1100: [W17] •MOXL [kg] K J108: (F1 ) [Wll] +SG:SES27«(J107-AL79) K L108: [Wll] •MOXL [kg] K 1108: (F1) [Wll] +SG:SES20*M1O7 K 1109: [ W17 J 'MP [kg] K J109: (FO) [Wll] @SUM(X79..AA79)+£SUM(AE79..AH79)+0SUM(AL79.. A079) K L109: [Wll] 'MP [kg] K Ml09: (F1 ) [Wll] <5SUM(C79.,F79)+gSUM( J79.. M79) +(ISUM(Q79 . .T79) K 1111: [ W17] •ML [kg] K Jill: (FO) [Wll] £SUM(J92..J103) K L1 1 1 : IW11] 'ML [kg] K ".111: (F 1 ) [Wll 1 0Sl'M(K92..M103) 409 K R88 [Will •SCSI K 090 [V5] ' TOTAL MISSION K P91 1*17) 'MA (kg) K Q91 (F2) (Wll) +J92+M92 K S91 [Will •ME (kg) K T91 (F2) (Wll) +J98+M98 K P92 (W17J •MG [kg] K Q92 (F2) (WLL] +J93+M93 K S92 (Wll) •USE (kg) K T92 (F2) [Wll] +J99+M99 K P93 1 W I 7 1 'MSS (kg) K Q93 (F2) [WU] + J94 + K94 K S93 I W11 J •MFS (kg) K T93 IF2) [Wll] +J100+M100 K P94 [ W 1 7) 'MRF (kg) K 09^ (F2) [Wll] +J95+M95+F111 K S94 I W 1 1 ] •KN (kg] K T94 (F2) (Wll] +J101+ !^101 K P9r, [W17J ' R0 X [kg] K Q95 ( F 2) (Wll] +J96+M96 + F103 K S-15 I W I 1 ) ,MFT [kg] K T95 < F 2) [Wll] +J102+V102 K S9fc |Wll) •KOXT [kg] K 796 ( F 2 ) (WU) +J103+"103 K P97 I W 1 7 J '"F Ikg] K Q97 B16 W17] 'SFB [W'/m2] C 1 6 till) +SG:SFS61*(SG:SFS15/C15)*2 Did Wll] +SG:SFS61"(SG:SFS15/D15)"2 El 6 Wll) +SG:SFS61*(SG:SFS15/E15)-2 F16 Wll] +SG:SFS61"(SG!SFS13/F15)*2 A16 W5] ' REFRIGERATION MASS PRELIMINARY CALCULATIONS B19 W17J ' Ox idi zer E1 9 Mil] • Fuel B20 W17 J 'HOX (W/m2) C20 Ft) (Wll] +$G:$FS50*$G:SFS21«$G:$F$65"t E 20 Wll] 'HF F 20 Ft) (Wll] +SG:SFS50"$G:SFS2O*$G!$F$63*t B21 W 17] ' ROXT (m/(kg)"l/3) C 2 1 Ft) (Wll] (0.75/(SG:SFS19"SG:SFS50*pPI))"(l/3) E21 Wll] 'RFT F 21 Ft) (Wll] (0.75/(SG:SFS17"SG:$F$t9'@PI))-(l/3) B2 2 W17] " Q1N (W/(kg)'2/3] C 2 2 Ft) [Wll] (C16*(l-SG:$FS5t)"$G:SFS2<.),@PI"C21-2 E2 2 Wll] " QIS F22 Ft) [Wll] (F16"(l-SG.SFS3t)"$G.'5FS23)"£PI"F21"2 B2 3 W17] ' 01. T| " " ] r2 j Ft) [WU] [C20"SG:SFS32)*t"ppi'C21"2 E2? Wll] goiT F23 Ft) [mi] (F20'SG:SFS31 i"t,@PI*F21"2 BUt W17] 'QCOOL [ " " ] L 2^ Ft) |kll] t$G:SFS50*(C22-C23) E2t Wll) 'QCOOL F24 f«) (Wll) +SG:SFStQ'(F22-F23) P:5 W17] MRPRE lkg'1/3] C25 Ft) [Ml] ( SG: SF$tt •C2t/SG: SFS52) E 2 Wll] '"RPRE F 2 5 Ft) |WU] I 5G: SFSt3'F2t/SG : SFS51 ) A.'T W5] ROCKET DESIGN' DETAILS (stage dependenc") Bi'fl W17] 'FEED SYSTEM LSED C 2 £ Wll] "pr» s s. r:e Wll] press. E2& Wll] "press. F 28 Wll) "press . B29 WIT] 'THRDES [kg) C 29 F0) [Wll] +SG:SFS66"C86 D29 F0) [Wll] +SG:$FS66*D86 E29 FO) [Wll] •SG:SFS66'E86 1 29 FU) [Wll] +SG:SFS66"F86 B 30 W17] 'TIMEB (sec] C 30 F2) [Wll] +C79*C13/C29 D30 F2) [Wll] eiF(D29=0,0,D79"D13/D29) E 30 F2) (Wll) @IF(E29»0,0,E79«E13/E29) F 30 F2) (Wll] 0IF(F29=O,O.F79"F13/F29) B31 W17] 'OPT. FEED SYSTEM C 31 Wll] @IF(C30<=(8130.1-C29--0.75867)," press."," pump") D31 Wll] g I F( D 29 = 0," n/a",§IF(D3O<»(013O.l«D29"-0.75867)." press. pump")) L : E 3 1 : Wll] piF(E29= 0," n/a",§1F(E30<»(6130.1'E29"-0.75867)," press. pump")) L : F 3 1 : Wll) 0IF(F29=O," n/a",@IF(F3O L B60 [W17] 'MF (kg) L C60 (F4) Will +C58+C59-C51/3 1. D60 (Ft) till) +D58+D59-D51/3 L E60 (FA) VI1) +E58+E59-E51/3 L F60 (F4) W11J •F58+F59-F51/3 L A62 [W5] MASS SUMMARY L C62 [Wll] '(Hohmann to highly eccentric Earth orbit) L B63 [ W17] •MA [kg] L C63 (F2) Will @IF(F7*" Yes "f/AND«F6« 1,$G SFS14*(C86 -C64) 0) L D63 (F2) Wll) flIF(F7 = "Yes"«ANDf'F6*2iSG SFS14*(D86 -D6'i) ,0) L E63 (F2) Wllj 0IF(F7«"Yes"WAND«F6=3 ,SG SFS1A"(E86 -E64) 0) I F63 (F2) Wll] §IF(F7«"Yes"»AND#F6"«,$G SFSH"(F86 -F6<0 ,0) L B6:SFS35"(D86-D6M+0 . k L Eo5 (F 2 ) Wll) + SG:SF$35ME86-E64)+0 .U 1. F 65 (F2 ) Wll] + SG:SFS35"(F86-F6<0+0 .k L [W17] ' MSS [kg] L ff)6 lF2) Wll] + SG: SFS60"(C86-C6<|) L D6b (F 2) Wll] •SG:SFS60"(D86-D64) L E6o (F2) Wit] + SG:SF$60'(E86-E6'<) L F66 (F2) Wll] + SG:SFS60,(F86-F6'<) L B67 [ W 1-1 •MRF 1 kg) L C67 (F2 ) Wll] gIF(F8="Yes",F25'C77"(2/3) ,0) L r»o 7 F2 ) Wll] |3I F( F8 = " Yes " i G25" D77 * (2/3).0) L E67 ( F 2) Wll] £IF(F8="Yes",H25"E77" ( 2/3 i,0) L F67 ( F2) k 1 1 ) £IF ( F 8 5 " Y e s ", I25-F77"(2/3) .0) L B66 [-17] 'v.KOX [kg] L •: ot ' F2 I » 1 1) piF(C8»"Yes '.C25"C70" (2/3) .0) L UbB 1 F 2 ) Wll] £ IF(C 8 » "Yes",D25"D78-(2/3).0) L Eo0 (F 2 J Wll] £ IF(C 8* " Y e s ",E 2 5 ' E 7 8 "(2/3) .0) L F6B (F2) Wll) @IF (C 8 = " Y e s ",F 25 " F 78 "(2/3) -0) L B7>, (H7J ' KE [kg] L C7Q (F2) Wll] +SG:SFS22"(C86-C6«) L P70 (F2) W1 1 ] + SG:SFS22"(D86-D6'i) L E "0 1F2) Wll) +SG:SF522"(E86-E64) L F 70 ( F 2 ) Wll] +SG:$F$22,(F66-F64) L B71 [W17] ' MSE [kg) L C 7 1 (F2) Wll] +SG:SF559*(C86-C6M L D71 (F2) Wll) •$G:SFS59'(D86-D64) IE 71 ( F2) Wll] +SG:$FS59"(E86-E6 L E73 (F2) (Xll] +SG:SFS47*(E86-E64) L F73 (F2) [Wll] +SG:SFS47«(F86-F64) L D74 [W17] 'KFT Ikg] L C 74 (F2) [Wll] +C37"(C86-C64) L D74 (F2) [Wll] +D37•(D86-D64) L 171, (F2) [Wll] +E37"(E86-E64) L f7>4 (F2) [Wll) +F37"(F86-F64) L B75 IW17J 'HOXT [kg] L C75 (F2) |h11) +C36•(C86-C64) L D75 (F2) |K11] +D36*(D86-D64) L H75 (F2) [Will +E36•(E86-E64) L F75 (F2) (Wll) +F36*(F86-F64) L 077 [W17) •«F [kg] I C 77 ' F 2) [Wll] + C60 L D77 ( F 2) [Wll] +D60 L E77 < F: I [Wll] + E60 L F77 ( F2 I [Wll) + F60 L B78 [W17) MOX [kg] L C 7 B (r 2 |W11) +C77»SC:SFS76 L [1-8 (F2) [Wll] +D77«SC:SFS76 L E78 • F 2 ) [W:1] +E77-SC:SFS76 L r-6 L: 116 [ W 1 7 •SFB tW/02) L:J16 [Wll +SG:SFS61*(SG:SFS15/J15)"2 L:K16 (Wll +SG:SFS61"(SG:SFS15/K15)"2 L: L16 (Wll +SG:SFS61"($G:SFS15/L15)"2 L: M16 (Wll +SG:SFS61'(SG:SFS15/H15)"2 L: H18 (N5] ' REFRIGERATION MASS PRELIMINARY CALCULATIONS L: 119 (W 17 Oxidizer L: L19 [Wll • Fuel L: 120 (W 1 7 'HOX [H/n2] L:J20 (F4) [Nil] +SG:$F$58'SG:$FS21*SG:$F$65- (Will ("20"S G : SFS31 )•4"gPI • M21 *2 L: I2<. 1 W 1 7 'QCOOL ( " " ] L :J 2 (F<<) |W11) +SG:$FS50MJ22-J23) L:L2<. (Wll QCOOL L : M2 160 [W17] ' MF 1 kg) J60 (F^) Wll) +J58+J59-J51/3 K60 (Ft) Will •K58+K59-K51/3 L60 (F'f) W11 ] +L58+L59-L51/3 M60 (FA) will +K58+M59-M51/3 H62 [W5] MASS SUMMARY 163 [W17] •MA [kg] J6 3 (F2) Wll) piF(M7="Yes"rtANDwM6«l,SG: SFS14 • (J86-J6 L M7 3 (F2) [Wll] £IF(J7«" No",SG:SFSA7,(M86-M6'(),M81'SH:SFS10) L 174 [VI 7] 'MFT [kg] L J74 (F2) [Wll ] 0IF(J7=" NO"iJ37'(J86~J64)IJ81'SH:SFS11) L K74 (F2) [Kill gIF(J7 =' N'o" ,K37'(K06-K6M ,K81'SH:SFS11) L L74 (F2) [Wll] 0IF(J7=" No" ,137" U86-L64),1.81 «SH:$F$11) , L M74 (F2) [Wll] £IF (J 7 « " No".M37*(K86-M6 L PIT) I K17 ] •SFB [W/m2] L Q16 [Wll] •SG:SFS61'(SG:SFS15/Q15)"2 L R16 ( K 1 1 J +SG:$FS61•($G:SFS15/R15)"2 L S16 (Kill +SG:SFS61'(SG:SFS15/S15)*2 L T16 [mi) + SG:SFS61«(SG!SFS15/T15 )"2 L 018 [W5] • REFRIGERATION MASS PRELIMINARY CALCULATIONS L P 1 9 IK17 J ' Oxidizer L S19 [Wll] • Fuel I P20 [ W 1 7 J 'HOX [W/m2] L Q20 (FA) [Wll] +$G:$FS58"SG:$FS21*SG:SFS65*4 L S 20 IWUj 'HF IT20 ( FA ) [Wll] +$G:$FS58*$G:$FS20*SG:SFS63"A I P 21 I W17] 'ROXT [m/(kg)"l/3] L Q21 (FA) [Wll] (0.75/(SG:SFS19"SG:SFS50*gPI))*(1/3) L S 2 1 1 W1 1| 1 RFT L T:I IFA) [Wll] (O.75/(SG:SFS17'SG:SFSA9*0PI))"(l/3) L F22 [ W 1 ~ ] QIS [W/(KG)"2/3] L Q-' 2 ( FA ) [Wll] (Q16 "I1-SG:SFS5A)• S G:SFS2A)*gPI• Q 21 "2 L S22 [Wll] 'QIS L T 22 ( FA ) [Wll] (T16"(1-SG:SFS3A)"SG:SFS23)"£PI"T21*2 L ?2 ) ( w 1 7 J 'QOIT [ " " ) L Q2 J < FA ) [Wll] (Q20,SG:SFS32),A'GPI'Q21'2 L 5 -: •>11) •QOIT L T2? ( FA ) [Will lT20"SG:SFS31)"A,gPI"T21"2 L ?2u ! U 1 7 ] 'QCOOL [ " " ] L Q2A (FA) [Wll] +SG:SFS50"(Q22-Q23) 1. ^ £ •* I W11 ] ' V,'CUOL 1 T 'FA I [Wll] +SG:SFSA9*(T22-T23) L P^FI !»>'.•] ' VRPRE [kg*1/3] L L P60 I W 1 7] •MF [kg] L Q60 (F<<) iwii] +Q58+Q59-Q51/3 L Rt>0 (F L: W1 5: (W17 ) ' RBS [AU] L:\15: (W11) +SB:SCS6 L:Y15: (W 11 ] +SB:SCS6 L:Z15: (W1| 1 +SB:SCS6 L:AA15: (W11) +SB:SCS6 L:Wlb: (K17J 'SFB [W/m2] L: X16: [Wll) +SG:SFS61,(SG:SFS15/X15)"2 L:\I6; [W11J +SG:SFS61*(SG:SFS15/Y15)*2 L:21 to: [Wll] + SG:SFS61•(SG:SFS15/2 15)*2 L: A.416 : [Will + SG:SFS61 *(SG:SFS15/AA15)* 2 L: V 18: [W5] ' REFRIGERATION' MASS PRELIMINARY CALCULATIONS L:W19: [ W17 ] ' Oxidizer L : Z 19: (Wll] ' Fuel L : U i Ci: IW 1 7] 'HOX IW / m 2] L:\2o: tF«) (Wll) +SG:SFS58*SG:SFS21"SG:SFS6'i* ii L:Z20: I In 11] ' HF LMOfl: IT") (Wll) + SG:SFS58*SG:SFS20"SG:SFS62" 1 I ] ' RFT L:\A21: lF<.) [Wll] (0.75/(SG:SFS16•SG:SFS49"@PI))*(1/3) t:k22: !<• 1Tj • Qlfi [W•, kg)" 2/3 ] L:X:S: i F<<) [ w 11 ] ( \ifj -(l-SG: SFS54 )*SG: SFS2A )*^PI"X21 *2 L:7..i: I'kll] • QIX I:i1 - 2: (FO [Wll] (AA16•(1-SG:SFS3i)'SG:SFS23)•@PI *AA21 " 2 L:w;i: 11.17] ' QOL'T [ " " ] L:\23: if>) [Wll] (X 20• S G : S F S 3 2) • <. *0PI * X21 " 2 L : 113 : ,w:i] 'QOIT 1 T A A 2 3: L:we 5: [W17| '^1 (kg] L : X85: (F2) [W11J gSl.'M(X63 . . X75) L : Y 85 : (F2) (Wll) gSLM(Y63..Y75) L:28S: (F2) (Wll] 0SUK(Z63..275) L:AA 8 5 i (F2) |W11] @SUM(AA63..AA75) L:W 86: (W17] 'WO [kg] L: X8t>: (F2) [Wll] +X79/( 1-X1« ) L: Y 86 : (F2) [Wll] ^IF(Y11,0,Y79/(1-Y14)) L: Z86: (F2) [Wll] eiF(Z14«l,O.Z79/(l-Z14)) L :A A d6 : (F2) [W11J @1F (AA14 =1,0, AA79/(1-AA1 L :AH 1i : [Wll) pIF(AH6=AtAE6/AH6,0) L : A[ 1 3: [ W 17) 'ISPl'SED [s] L: A t' 1 3: (Wllj +SG:SFS38 1 : AF 1 3: (Wll] +SG:SFS 38 L : AG 1 3: iWll] +$G:SFS30 L : A H 1 3 : [Wllj +SG:SFS30 L:AD I 4: (W17) 'M R A T10 L: A E 1 AD 16 17 'SFB |K/ra2] AE16 11 +SG:SFS61" (SG:SFS15/AE15) "2 AFlb 1 1 +S G : S F5 61 "(SG:SFS15/AF15)*2 AG16 11 +SG:SFS61'(SG:SFS15/AG15)*2 AH16 11 +SG:SFS61'(SG:SFS15/AH15)"2 AC16 5) • REFRIGERATION MASS PRELIMINARY CALCULATIONS AD1 9 I7 ' Oxidizer AG 19 ]1 ' Fue 1 A D2 0 17 'HOX (K/m2 J AE20 m [W11 ] •SG:$FS58 ,SG:$FS21*SG:SFS6 L AD38 : IW17 ) ' FSC L A E 38 : (F 4 (Wll) gIF(AE28»"press.",SG:SFS29*SC:SFS68/500,$G:SFS30*gEXP(0.00075 9 SC:SFS68) L AF38: lF4 [Wll] gIF(AF20*"press.SG:SFS29-$C:SFS68/500,SG:SFS30'0EXP(O.00075 9 SC :SFS68) L A G 38 : (F4 (Wll) 01F(AG28»"press.SG:SFS29*SC:SFS68/500,SG:SFS30*0EXP(O.00075 9-SC:SFS68) L A H 38 : (F4 (Wll) glF(AH28«"press.",SG:SFS29*SC:SFS68/500,SG:SFS3O«0EXP(O.00075 9*SC:SFS68) L AC.O: [V5 ' INTERMEDIATE MASS CALCS I ADftl : (W1 ) 'COEFSUH L A E A l : (Fft (Wll) +SG:SFS35+SG:SFS^7+AE37+AE36+SG:SFS59+SG:SFS60+AE38+SG:SFS22 L AFftl : (Fft [Wll] +SG;SFS35+SG:SFSft7+AF37+AF36+SG:SFS59+SG:SFS60+AF38*SG:SFS22 L AG ft 1 : (Fft [Wll] +SG:SFS35+SG:SFSA7+AG37+AG36+SG:SFS59+SG:SFS60+AG38+SG:SFS22 L AHftl : (Fft (Wll) +SG:SFS35 + SG:SFS't7 + AH37 + AH36 + SG:SFS59 + SG:SFS60 + AH38 + SG:SFS22 L AD4 2: [W1 ] 'COEFTL L AEfti: (Fft [Wll] (JIF(AH7 = "YES"«A!iD»AH6 = l,SG:SFSl'i + AEftl.AE L AD71 (U 17 ] ' MSE [kg] L AE71 (F2) WU| +$G:SF$59*(AE86-AE6 : A K 16 1w17] 'SFB [W/m2] :A L 1 6 [Wll] + SG: SFS61"(SG:SFS15/AL15)"2 :AM6 [Wll] +SG:5FS61"(SG:$FS13/AK13)"2 :A N16 [Wll] +SG!SFS61*(SG:SFS15/AN13)*2 : AO 16 [Ml] +SG:SFS61»(SG:SFS15/A015)*2 r A J1 6 [W5] REFRIGERATION KASS PRELIMINARY CALCULATIONS : AK19 [W17J ' Oxidizer :A S19 [Wll] ' Fue 1 : A K 20 [ W17 ] 'HOX [W'/m2] : AL20 (FA) [Wll] •$G:$FS58"$G:SF$21'$G:SFS6A*A : AS 20 [Wll] 'HF : AO20 (FA) [Wll] +SG:SFS58'$G:$F$20*$G:$F$62'A : A K 2 1 [W17] 'ROXT fm/(kg)* 1/3] : A L 2 1 (fa) [Wll] (O^S/lSGlSFSie'SG.-SFSSO'pPI))- (1/3) ;AV21 [Wll] 'PFT : A0 2 1 (fa) [Wll] (0.75/($G:SFS16*SG:$FSA9"(aPI))-(l/3) : AK 2 IW17] 'q1x [W/(kg)"2/3J : al2; iFA) [Wll] (AL16"(l-SG:SFS5A)*SG:SFS2A)*pPl"AL21*2 L : AS 22 i w 1 1 ] •QIX L : AO 22 IF 4 ) [Will (A016«(l-SG:SFS3A)«SG:SFS23)'gP] VA021 * 2 L :AK2j i k;:] •qoit : AL 2 3 ( r a ) [Wll] ( AL20"SG: SFS32)*A'@PI\AL21 "2 : av2t i W 1 1 ] • QOL T A 12 3 IF-) [Wll] ( A020* SG :5 F S 31 ) * 4 •A021 • 2 : A K 2a [W17j 'QCOOL [ " " ] :A L i <• if * > I W 1 1 ] + SG:SFS 501(AL22-AL23) : AN 2 a [Wll] ' ijC 00L : A02a (fa; [Wll] +SG:SFSA9"(A022-A023) Ah 2 "> l»17] "IRPRE (kg" 1/7] : <1.25 ( FA ) (Wll] ($G:SFSAA "AL2A/SG:SF5 5 2 ) \S 2 r. 'VRPRE : A02! ( FA ) (nil] (SG:SFSA3"A02A/SG:SFS51) : A J 27 iw1] ROCKET DESIGN DETAILS (stage dependence) :AK28 [W17] 'FEED S\STE* ISED i w 1 1] pomp A v 2 •? [Wll] "pump [hi] " pump A028 [wll] pump \Kj<) 1 w 1 7 ] •THRDES [kg] L : A I^ i f0) [Wll] + S G:S FS66••AL86 l:a*29 (FO) [Wll] +SG:SFS66" L:AS 2 i (FO) [Wll] + SG:SFS66• ' A S8 6 i.: AO29 F 0 ) [Wll] +SG:5FS66'A086 L : >0 IF 2 ) [ W 17 ] 'TI"EB [sec] L :A L 30 (F2) [Wll] +AL79"AL13/AL29 L 30 (F2) (Wll] ^IF(AM29"0,0,A^79*AM13/AM29) L:AS 30 (f2) [wll] 0IF(AK29= O,O, A(J7 9*AN13/AN29) L:A030 (F2) [Wll] eHF(A029= 0,0,A079«A013/A029) L AK 3 1 n 'OPT FEED SYSTEM l AL 3 1 [Wll] pIF(AL30 < =(8130. 1 AL29' -0.75867)," press."," pump") L A f. 3 1 [Wll] 0IF(AM29= O, " n/a' i@IF(AM30<=(8130.1*AK29*-0.75867)," pres s pump")) L : A S;«1 : [Wll] pIF(AN29»0," n/a",gIF(AN30<«(8130.l*AN29"-0.73867 ) ,• pres s . " , " pump")) L: A031 [Wll] @IF(A029=0," n/a",gIF(A030<=(8130.1•A029--0.75867),' pres pump")) I: AK 3 2 : (W17 ] • PT [MPa ) L:A L 3 2 : (F4) (Wll] gIF(AL28"" Pl'HP", 50"0 . 101325/14 .696 ,1. 5*SC : SFS68 • 0.101325/14 .6 96) L:AH32: c F ) [Wll) gIF(AH28'="Pl'HP" ,50*0.101325/14.696, 1. 5 • $C : SFS60 •0.101325/14 , 6 9ft) L: A S3 2 : (F4) [Wll] CIIF(AN28=" PUMP" ,50*0.101325/14. 696, 1. 5" SC: SFS68 •0.101325/ 1 4 , 6 96) lm032: (F4) [Wll] gIF(A028«"PUMP",50*0.101325/14.696,1.5*SC:SFS68*0.101325/14.6 96) l:ak 33: [W17] 'PINDEX L: AL33 : (F0) [Wll] gIFCAL32>"SD:SIS71,7,glF(AL32>=SD:SHS71,6,gIF(AL32>=SD:SGS71 , 1,g IF ( A L 3 2 > =S D: SFS71 , 4,gIF(AL32>«SD:SES71 , 3,$D:SJS61 )) ) )) L:33: (FO) [Wll] gIF(AK32>= $D:SIS71,7,g1F(AM32> = SD:SHS71,6,$ IF(AH32>»SD:SGS71, 5,plF 5,0IF L: AOr,T: (F2) [Will 01F f A077= 0 ,0 ,p IF (A08-" Ye s "«ASD«SC :S GS 74=" .So "«AND« S C :SES74 = " N 0 "••••ASD-'.-SC: SDS65=SC : S DS7, AQ25" ( A07 7+SCS127) * (2/3) ,0IF( AOS" "Yes" ,AQ25"A077" (2/3) ,0 ) J) L:AK 68: [W17] '^ROX [kg] L:Ue>8: (F2) [Wll] 0IF ( AL8= "Ye s "»ANDWSC :S G S 75=" No "»AND« SC :S ES75= "No" »AND»$C: SDS 6 6==.-:b&s75.AL25*(AL78+ Q119)-(2/3),0IF(AL8""Yes",AL25"AL70-(2/3),O)) L:^v68: ;F2) [Wll] 0IF(AM78=0,0,0IF(AL8»"Yes"«AND«SC:SGS75•"No"«AND«SC:SES75="No ASD»SC :SDS66=SC : SDS75 ,AL25 '(AK7B + Q119 ) * ( 2/3 ) ,01F (AL8- " Ye5" .AL25 • AK78" (2/3) ,0 ) ) ) L :A S 60: (F2) [Wll] 0IF(AN78= 0,0,01F(AL8 ="Yes"tfAND«$C:SGS75»"No"#AND0$C:SES75«"No "»»ASD«SC: SDS66= SC : SDS75 ,AM25 • (AN78+ R119) " ( 2/3 ) ,0IF(AL8•" Yes " ,AM25" AN78* ( 2/3) ,0) ) ) L: A 06 8 : (F2) [Wll] 01F ( A078«0 ,0 ,01F (AL8« " Yes " tfANDtfSC :S GS 75 = "No "«AND«SC: SE$75« " No "KASD.VSC: SDS66*SC : SDS75, AN25* (A078+ S119)- (2/3) ,0IF(AL8«"Yes" ,AN25' AO70" (2/3) ,0)) ) L :A K 70 : [W17] 'ME (kg) L:AL70: (F2j [Wll] + SG:SFS22*(AL86-AL64 ) L:A"70: 1F2) [Wll) +SG:SFS22"(AM86-AM64 ) L: A S 70 : (F2 J [Wll] +SG:SFS22•(AS86-AN64 ) L:A0 7 0: lF 2) [Wll) + SG: SFS22 *(A086-A064 ) 452 L AK71 [W17] 'V.SE i kg 1 L AL71 (F2) will + $G:SFS59"(AL86-AL6 L F1 15 (FA) [Mil] +F78 L E1 1 6 [ W 17] "LEG 5 L CI 16 (FA) [Wll] +J78 L D116 (FA) [Nil] +K78 L E116 (FA) [Wll] +L78 L F116 (FA) [Kll) +H78 L B117 [W17] "LEG L E1I 'V [Will 'MASS LSED L F11".77 1. PI [WIT] "LEG A L C 1 25 (FA i |W11] +Q77 L D! 25 ( FA ) [Wll] + R77 L E 1 25 (FA) (Wll) +S77 L F;:V (fA ) [Wll] +T77 L B: 27 : •» : 7 J "TOTAL L C: (FA ; [Wll] gSlM(C123..F125) I_ L'L 27 [Wll] > L El 27 IWU] '"ASS USED L F127 (FA ) [Wll] gIF($AAS8="Yes",C127,0) H89 (W 5 ] 'TOTAL MASS SUMMARY J89 [Wll] +SCS1 H91 t W 5 ] *• PRE-MARS K 91 (Wll] 1 POST-MARS 192 Ih17] 'MA [kg] J92 (FO) [Wll] @Sl'M(X63. .AA63)+@SUM(AE63. .AH63)+gSUM(AL63. .A063) L92 [Wll] 'MA [kg) M92 (Fl) [WU] @StiM (C63. .F63)+@SUM(J63..H63)+@SUM(Q63. .T63) 193 IW17] 1KG [kg] J93 (FO) [Wll] @SUM(X65. .AA65)+gSUM(AE63. .AH65)+f>SUM (AL65. .A065) L93 [Wll] 'MG [kg] M93 (Fl) [Wll] @SUM(C65..F65)+0SUM(J65..M65)+@SUM(Q65..T65) 194 IW17] 'MSS [kg] J9d IFO) [Wll] @SIM(X66. .AA66)+0SL'M(AE66. . AH66) (AL66. .A066) L9i. [Wll] "MSS [kg] m9" (Fl) [Wll] @SIM(C66. .F66 )••{ISUM ( J66. .M66)+@SUM(Q66. .T66) 195 (W17) 'MRF [kg] J95 (FO) (Wll) gSlM (X67.,AA67)+pSUM(AE67..AH67)+pSUK (AL67..A067) L95 IW11] 'MRF [kg] "95 (Fl) [Wll) @511(067. .F67)[email protected](J67. .M67)+@Sl'M(Q67. ,T67) I96 [WIT] 'MROX [kg] 395 (Fu) [Wll] pS L ^. ( X 60 . .A A6fl )+ @S I'M ( AE68 . .AH68 )t@Sl M ( AL68 . .A068) L9i ['.11] 'MROX [kg] '90 itl) [Wll) @SIM(C68. .F68)[email protected]( J68. .M68 ) t@SLM(Q68. .T68) 198 |W17) 'ME [kg] J 98 (FO I [Wll] @SIM(X70. .AA70)+£SIM(AE70. .AH70)+0SIM(AL70. .A070) 1.98 (Wll] 'VE II- g} "'j 8 (Fl) [Wll) @SIM(C70. .F70)+gSUM(J70. .M7 0 )+PSUM (Q70 . .T70) ;9 9 [W17j ' MSE (kg] .19 9 < I0 J |W1>) @3','v(\71 . .AA71 )+@Sl:M(AE71 . .A H7 1 )+ @St'M ( AL7 1. . A07 1) L99 (Wll) '\SE [kg] m 99 ( F 1 I [Wll] pSI.'M(C71 F71 )+gSlM(J71 . ,M71 )+gSLM(Q71 . .T71) L 1100 1 W 17 ] 'NFS (kg) L JlOO (FO) [Wll| @SIM(X72. .AA-72)+0Sl:M(AE72 ..AH72)*0SUM(AL72. .A072 ) L LlOO 1 »! 1 ] ' NFS [kg] L MOO (Fl) [Wll] 0Sl'M(C72. .F 7 2) •» @S L M( J7 2 . .M72)+@Sl'M(Q72. .T72) L i:o i 1 W 17 J ' N.N [kg] L J1 01 (FO) IV. 11J @SIM(X7 3. .AA73)+@SUM(AE73 ..AH73)+0SUM(AL7 3.. A073) L L 1 01 [Wll] ' MS [kg] L "101 (Fl ) [Wll] @Sl'M (C73. •F73)+gSUM(J73..M73)+@SUM(Q73.,T73) L 1 102 [ W 17 ] ' MFT [kg] L J102 (FO) (Wll] gSUM(X74. .AA74)+GSUM(AE7< I ..AH7*i)+@SUM(AL74. .A07 4 ) L Lin; [Wll] ' MFT [kg] L "102 (Fl ) [Wll] @SUM(C7<.. .F74 )+ @Sl'M (J74 , .M7<.)+@SUM(q7<., .T7^) L 1103 [W17] ' M0XT [kg] L J1 03 (FO) [Wll] @SUM(X75. .AA75)+@SUM(AE75 ..AH75)+@SUK(AL75.. A073) L L 10 3 [Wll] 'MOXT (kg] L "103 (Fl ) [Wll] pSL'M (C75 . .F75)+ gSl :M( J75..M75)+0SUM(Q75..T75) L I105 IW 17 ] '«F [kg] L 310 5 (FO) [Wll] @SUM(X77. .AA77)+gSUK(AE77 ..AH77)+gSUM(AL77.. A077) L L 105 [Wll] ' KF [kg] L "105 (Fl) [Wll] pSIM(C77. .F77)+@SUM(J77.. M77)+0SUM(Q77..T77) -Fl27 L I106 ( W 1 7 ] 'NFL [kg] L J1 06 (Fl) [Wll] +SG:SFS 25 •(J105-AL77) L LlOo (Wll) ' MFL [kg] r-r-r-r-r-r-r-r-r-t-r-r-r-r-t-r-r-* ' l/l os 457 L R88 [WllJ +SCS 1 1 090 [W5] • TOTAL MISSION L P91 [ K17 ] •r-A lkg] L Q91 (F2) [WU] +J92+M92 L S91 [»ill] •WH Ikgl L T91 (F2) [Wll] +J96+M90 L P92 [ W 17 ] 'MG [kg] L Q92 (F2) [Wll] +J93tM93 L S92 (Wll] 1MSE [kg] L TO 2 (F2) [Wll] •J99+M99 L P93 [ V. 171 'MSS [kg] L g REFERENCES 1. Ramohalli, K., Dowler, W., French, J., and Ash, R., "Some Aspects of Space Propulsion with Extraterrestrial Resources," Journal of Spacecraft and Rockets, Vol. 24, No. 3, May- June 1987, pp. 236-244. 2. Zubrin, R., Baker, D., and Gwynne, O., "Mars Direct: A Simple, Robust, and Cost Effective Architecture for the Space Exploration Initiative," 29th Aerospace Sciences Meeting, AIAA 91-0326, Reno, NV, Jan. 1991. 3. 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Preiss, B., Pan, T., and Ramohalli, K., "Further Applications of a Figure-of-Merit in Space Missions," 27th Joint Propulsion Conference, ALAA 91-2330, Sacramento, CA, June 1991. 0 z: • V»w - CO w C) O (h • «>H «) • V)H O <7 1 «»9 'T • 9 NN • U* h- co - (&« CO «-t CN CN V>«H(N CO 03 H H M A - O H V>Q • U N wr> a> u« • CO «] u* nhiOJS • CO U« (D XI m HM17U CN HH V>U CO ** AJjCQ r »-• m>n « H Q «v> • ^>4 O - CO - wt«> a O O w r- • 0 ts U* - M «-4 » fau - CO »-4 H O H *• t» H-« N r- to &e-» ^ CD - CO Q CO CO •V)QO co as 9 CiO^v) »-« +»