Orbital Fueling Architectures Leveraging Commercial Launch Vehicles for More Affordable Human Exploration

Home , Aerocapture, Apollo Docking Mechanism, Exploration Upper Stage, Hypergolic propellant, Isogrid, Launch window

ORBITAL FUELING ARCHITECTURES LEVERAGING COMMERCIAL LAUNCH VEHICLES FOR MORE AFFORDABLE HUMAN EXPLORATION

DANIEL J TIFFIN

Submitted in partial fulfillment of the requirements for the degree of: Master of Science

Department of Mechanical and Aerospace Engineering

CASE WESTERN RESERVE UNIVERSITY

January, 2020 CASE WESTERN RESERVE UNIVERSITY SCHOOL OF GRADUATE STUDIES

We hereby approve the thesis of DANIEL JOSEPH TIFFIN

Candidate for the degree of Master of Science*.

Committee Chair Paul Barnhart, PhD

Committee Member Sunniva Collins, PhD

Committee Member Yasuhiro Kamotani, PhD

Date of Defense 21 November, 2019

*We also certify that written approval has been obtained for any proprietary material contained therein.

Table of Contents

List of Tables...... 5 List of Figures ...... 6 List of Abbreviations ...... 8 1. Introduction and Background...... 14 1.1 Human Exploration Campaigns ...... 21 1.1.1. Previous Mars Architectures ...... 21 1.1.2. Latest Mars Architecture ...... 26 1.1.3. Fueling Architectures ...... 27 1.1.4. Logistics and Tracking ...... 29 2. Methods ...... 32 2.1. The Systems Engineering Process ...... 32 2.1.1. Systems Analysis and Control ...... 33 2.1.2. Systems Analysis in Practice ...... 34 2.1.3. Architectures ...... 35 2.1.4. Architecture Definition ...... 36 2.2. Propellant Tracking Theory ...... 40 2.2.1. Capabilities and Rationale for Propellant Tracking ...... 40 2.2.2. Program Logic and Structure ...... 41 2.2.3. Program Verification ...... 44 2.3. Campaign Analysis ...... 46 2.3.1. Mars Hybrid Propulsion System Campaign ...... 47 2.3.2. Lunar Exploration Campaign ...... 64 2.4. Useful Concepts ...... 69 3. Results and Discussion ...... 73 3.1. Mars Hybrid Propulsion System Campaign Trades ...... 73 3.1.1. Hypergolic Chemical Propellant Alternative ...... 73 3.1.2. Latitude Sensitivity...... 75 3.1.3. Tanker Design Sensitivity ...... 78 3.1.4. Lander Trade ...... 80 3.1.5. Tanker Thermal Control System Trade ...... 81 3.1.6. Propellant Transfer Rate Trade ...... 82 3.2. Lunar ...... 84

3.2.1. Fueling Architectures ...... 84 3.2.2. Element Designs ...... 84 3.2.3. Performance ...... 94 4. Conclusions and Future Work ...... 105 Appendix ...... 110 References ...... 116

List of Tables

Table 1. Previously Proposed Mars Architecture Breakdown ...... 23

Table 2. Potential Risks ...... 31

Table 3. Summary of the Spacecraft Design and Sizing Process ...... 39

Table 4. Hybrid Propulsion System Mass Breakdown ...... 47

Table 5. Baseline HPS Summary ...... 48

Table 6. Launch Vehicle Capability Assumptions ...... 52

Table 7. Passive and Active TCS Hybrid Tanker ...... 54

Table 8. Mars Crew Campaign: Storable HPS ...... 56

Table 9. Storable HPS Mass Breakdown ...... 56

Table 10. Storable HPS Key Information ...... 57

Table 11. Passive CFM Hybrid Tanker ...... 63

Table 12. Propellant Need Assumptions: Human Lunar Lander ...... 68

Table 13. Launch vehicle capability assumptions ...... 68

Table 14. HPS Fueling Window...... 83

Table 15. HPS Order of Magnitude Fluid Transfer Rate Requirements ...... 83

Table 16. Propellant Thermophysical Data...... 85

Table 17. Propellant Delivered to NRHO ...... 95

List of Figures

Figure 1. Close-up of Progress Spacecraft docking to ISS ...... 15

Figure 2. Restore-L Conceptual Rendering ...... 18

Figure 3. Explanation of currently planned SLS Block configurations...... 24

Figure 4. The Systems Engineering Process ...... 32

Figure 5. PropTracker Basic Code Structure ...... 42

Figure 6. PropTracker Sample of Graphical Output ...... 44

Figure 7. Hybrid Propulsion System ...... 47

Figure 8. Mars Campaign to be used as a basis of comparison ...... 52

Figure 9. Visualization of NRHO proposed for Gateway ...... 53

Figure 10. Storable HPS Model...... 56

Figure 11. Storable HPS: Crew-Only Campaign Full Factorial...... 74

Figure 12. HPS Campaign: Latitude Sensitivity on Fueling for 15 t CLV Tankers ...... 77

Figure 13. HIAD Lander Campaign: Tanker Design Sensitivity ...... 78

Figure 14. Lander Mass Sensitivity ...... 80

Figure 15. Cryogen Thermal Control Trade ...... 82

Figure 16. Expendable Tanker Mass Fraction Delta: Total Lunar Campaign Refueling .. 96

Figure 17. Refueling Flights: 7.1 t Tanker Inert Mass ...... 97

Figure 18. 50 Missions: No. of Stages & Elements Expended: 7.1 t Inert Tanker ...... 98

Figure 19.Passive Fast v. Active Slow Fueling Architecture: Delivered Propellant ...... 100

Figure 20. Propellant Delivered to NRHO: Upper Stage Stays in NRHO ...... 101

Figure 21. Propellant Delivered to NRHO: SEP Tug; Upper Stage Stays in NRHO ..... 102

Figure 22. Propellant Delivered to NRHO: SEP Tug ...... 104

Figure 23.Refueling Architecture Comparison: Propellant Delivered to NRHO ...... 105

A- 1. Human Landing System: Baseline 3-Element Architecture ...... 111

A- 2. Refueling Element (CLV Upper Stage or Tanker) ...... 112

A- 3. Refueling Element (Either CLV Upper Stage or Tanker) + Reusable Bus ...... 113

A- 4. SEP Tug + CLV Upper Stage: Upper Stage Ends in NRHO ...... 114

A- 5. SEP Tug + CLV Upper Stage: Upper Stage Returns to LEO ...... 115

List of Symbols

휔 Angular velocity (rad/s) 3 휌푖 Insulation density (kg/m ) 휇 Standard gravitational parameter (km3/s2) 푐 Constants of integration 퐶 Jacobi constant 퐹 Heat flux (W/m2) 푔 Gravitational acceleration (m/s2) 퐺 Universal Gravitational constant (km3/kg/s2) 푚 Mass of a planetoid (kg) 푀 Total system mass (kg)

푀푓 Spacecraft mass fraction 휆 Oscillation frequency, in-plane (rad/s) 휈 Oscillation frequency, out-of-plane (rad/s) 휙 Phase angle, in-plane (rad) 휓 Phase angle, out-of-plane (rad)

퐴푦 Oscillation amplitude, in-plane (km)

퐴푧 Oscillation amplitude, out-of-plane motion (km) 퐻 Heat of vaporization of propellant (kJ/kg) 퐾 Thermal conduction coefficient (W/m) 푟 Radius (km) 푅% Boiloff rate as a percentage of total mass (%/day), see Equation 14

푇 푇푒 − 푇푏 , See Equation 13

푇푒 Tank external temperature (K)

푇푏 Tank internal (boiling) temperature (K) Δ푇 External/internal surface temperature difference (K), see Equation 15 푡 Time (s) Ų Sum of the centrifugal and gravitational potentials 푣 Velocity (km/s)

푋 Insulation thickness (m) 푥, 푦, 푧 Cartesian position coordinates (km) 푥̇ , 푦̇ , 푧̇ First time derivative of the Cartesian position coordinates (km/s) 푥̈ , 푦̈ , 푧̈ Second time derivative of the Cartesian position coordinates (km/s2)

List of Abbreviations

AE: ascent element

AR&D: autonomous rendezvous and docking 18

BLT: ballistic lunar transfers 51

BOL: beginning of life 48

CBH: common bulkhead 82

CFM: cryogenic fluid management 34

CLV: commercial launch vehicles 17

COTS: commercial off-the-shelf 84

CP: chemical propulsion 19

CRTBP: circularly restricted three body problem 68

DAE: descent-ascent element

DARPA: Defense Advanced Research Projects Agency 14

DE: descent element

DLO: densified liquid oxygen 83

DRA: Design Reference Architecture 19

DST: Deep Space Transit 48

ECLSS: environmental control and life support systems 29

EMC: Evolvable Mars Campaign 19

EOR: Earth orbit rendezvous 12

EUS: Exploration Upper Stage 19

GTO: geosynchronous transfer orbit 65

HET: Hall-effect thruster 57

HIAD: hypersonic inflatable aerodynamic decelerator 53

HPS: Hybrid Propulsion System 24 iSR: in-space refueling 34 iSSA: in-space servicing and assembly 34

JPL: Jet Propulsion Laboratory 19

JSC: Johnson Space Center 14

LOR: lunar orbit rendezvous 12

Mid-L/D: mid-range lift to drag ratio 53

MLI: multi-layer insulation 60

MOI: Mars orbit insertion 20

MPS: main propulsion system 60

MTV: Mars transit vehicle 23

NASA: National Aeronautics and Space Administration 12

NEA: near Earth asteroids 17

NRHO: near-rectilinear halo orbit 24

NTP: nuclear thermal propulsion 19

ORS: Orbital Refueling System 14

ORU: orbital replacement unit 15

PAF: payload adapter fitting 18

RCS: reaction control system 85

RE: refueling element

RRM: Robotic Refueling Mission 15

SEP: solar electric propulsion 41

SLS: Space Launch System 22

SoS: Systems of systems 33

SpaceX: Space Exploration Technologies Corporation 23

STR: space technology roadmap 35

TCS: thermal control system 87

TE: transfer element

TLI: trans lunar injection 48

TM: thermal mechanical pump 14

TMI: trans Mars injection 20

TRL: Technology Readiness Level 34

ZBO: zero boiloff 15

Orbital Fueling Architectures Leveraging Commercial

Launch Vehicles for More Affordable Human Exploration

Abstract

DANIEL J TIFFIN

To fuel transportation systems, there exists an opportunity to reduce launch costs by an order of magnitude by launching the necessary propellant on existing commercial launch vehicles (CLVs). This research analyzed various architectures that deliver propellant to near-rectilinear halo orbit (NRHO). An automated tool was developed and utilized to rapidly trade architectures. First-order results indicate many feasible architecture options exist for commercially launched propellant. Active cryogenic fluid management (CFM) tankers were shown to have negligible improvements over passive tankers that rendezvous with a reusable (active CFM) bus. CLV long-duration upper stages deliver more propellant than ZBO tankers if, on average, tanker inert mass is greater than 51% of the CLV usable payload. “Topping-off” long-duration upper stages with propellant in LEO permits a mean of 13 metric tons per launch delivered to NRHO.

Reusable tugs were shown to increase delivered propellant per launch by 180% on average.

1. Introduction and Background

Space exploration has a long history of countless technology demonstrations and advances that allow humans to fly farther, faster, and for longer. Remarkable achievements have been made in virtually every facet of space travel—from the rovers put on Mars to the astronauts who walked the moon, insurmountable challenges and missions we once thought impossible were achieved. After more than half a century of learning how to fly space vehicles, one capability has remained particularly elusive: autonomous in-space propellant transfer.

Transferring fluids between two spacecraft can be complicated. The National

Aeronautics and Space Administration (NASA) knew this even back in the days of the

Apollo program. In fact, one of the principle reasons John Houbolt’s famous single- launch architecture, the lunar orbit rendezvous (LOR), beat out the less risky Earth orbit rendezvous (EOR) approach and direct-ascent approach (championed by Wernher von

Braun) was because LOR did not require additional Saturn V launches to fuel a more massive, monolithic spacecraft [1]. By 1962, NASA was under immense pressure to choose an architecture; fueling a vehicle in space had never been attempted, and fewer launches meant lower cost. In-space fueling appeared to be one more hurdle that could have cost the United States unacceptable schedule slips or altogether mission failure.

As history would have it, LOR turned out to be the right approach for Apollo, but today, we not only want to explore farther destinations and operate in space longer, but the launch vehicles available now are dwarfed by the capabilities of the gargantuan

Saturn V. The desire to launch more payloads, and the need to launch in smaller chunks,

14 not surprisingly, makes in-space propellant transfer practically unavoidable. Countless studies have cited orbital fueling as a critical need to enable humans to continue and advance space exploration [2]–[5].

One might expect in-space fueling to be routine by now. In fact, only a handful of vehicles have transferred fluids in space. The first being the Russian made and operated

Progress spacecraft, which can pass hypergolic propellant and water to the International

Space Station (ISS) through ports in the docking ring. Progress M1 has zero water tanks and eight propellant tanks (four for fuel and four for oxidizer), which can carry up to

1,740 kg of propellant, depending on how much mass is allocated to cargo. The M variant of Progress has six tanks—two fuel, two oxidizer, and two water tanks. Progress has

Figure 1. Close-up of Progress Spacecraft docking to ISS Source: http://www.nasa.gov/sites/default/files/iss040e070868.jpg

15 routinely transferred propellants to ISS since 1978 [6]. Transfers have generally been successful for Progress, but at least one leak in the water supply system has occurred [7],

[8]. Few additional details on Progress’s hardware is publicly available. Fortunately, knowledge of the general configuration is sufficient for the purposes of this investigation.

The Orbital Refueling System (ORS), built by NASA Johnson Space Center

(JSC), was launched in 1984. ORS demonstrated six back and forth transfers of 142 kg of hypergolic propellant. The vehicle used in-tank flexible bladders for liquid acquisition, much like Progress [6]. Bladders are effective for storable propellants because they keep the pressurant separate from the propellant. This prevents vapor ingestion during transfer.

Extremely low temperatures currently prohibit bladders from being implemented in cryogenic tanks. Other clever liquid acquisition techniques and microgravity tank settling techniques have been proposed, however.

Cryogen orbital transfer has only been successful once to date. The Super-Fluid

On-Orbit Transfer (SHOOT) flight demonstration, flown in 1993, successfully passed super-fluid helium between two dewars. One of which used a screen channel, the other had a vane device for liquid acquisition. A thermal mechanical pump (TM) transferred the helium at a maximum rate of 720 l/hr from the vane side and a maximum of 385 l/hr from the screen channel side [6]. Although the demonstration was successful, SHOOT relied on unique properties of super-fluid helium for the TM to work. The experiment’s applicability to other cryogens is limited.

Orbital Express, a Defense Advanced Research Projects Agency (DARPA)-

NASA joint mission, demonstrated the feasibility and utility of fueling a satellite designed to be refueled. (Another mission accomplishment was transferring an orbital

16 replacement unit (ORU) to the visiting vehicle using a robotic manipulator. The ORU is effectively a box that can accommodate new batteries, a flight computer, instruments, etc.) The servicing vehicle autonomously rendezvoused and berthed with the visiting satellite and successfully transferred hydrazine via two methods: ullage recompression as well as a fluid pump [9].

There are currently thousands of satellites in orbit, and only two were designed to be serviceable—the Hubble Space Telescope and ISS [10]. To address the immediate need for orbital refueling, NASA has focused on maturing the capability to refuel satellites not specifically designed to be serviceable. One series of missions to demonstrate this capability was NASA’s Robotic Refueling Mission (RRM). RRM1 successfully demonstrated ethanol propellant transfer on ISS. RRM3 demonstrated a zero boiloff (ZBO) thermal control system for four consecutive months on ISS. The goal was to first prove ZBO for long durations, then attempt cryogen transfer, but a cryocooler failed before cryogen transfer could even be attempted [11]. Although the history of in space fluid transfer is limited , the capability is gradually maturing and becoming more feasible.

Even in the next several years, there is strong evidence that orbital fueling will continue to advance. Another mission similar to RRM, called Restore-L, will service and refuel government owned Landsat-7 in 2020. Restore-L will rendezvous, grip Landsat-7, cut through MLI, and attach a hose to pressure feed hydrazine, thereby extending the lifetime of the satellite. [3]. Additionally, a small startup company called Orbit Fab demonstrated liquid water transfer on ISS less than a year after the company was

17 established [12]. Orbit Fab also recently secured over $3 million in funding to provide satellite refueling services to a wider market [13].

Figure 2. Restore-L Conceptual Rendering

Source: https://sspd.gsfc.nasa.gov/images/Restore_Image_300ppi_021web.jpg

Orbit Fab has ambitions to create a market for in-space fueling, but they are not alone in this endeavor. SpaceX and NASA recently announced a partnership to mature orbital refueling technology—a collaboration that could enable both organizations’ goals of lunar and Mars exploration [14]. From NASA, to industry, to military, to startup companies, the need for in-space fueling is not only apparent, but widespread.

Historically, the concept of fueling a vehicle in space used to be enough of a potential risk to bar missions involving fueling from serious consideration. We have learned from both ISS and numerous demonstration missions that fueling can be done safely and routinely in space. The time has come to increasingly rely on fueling as a mission enabling capability. Space exploration began as a series of “one-off” missions

18 involving billions of dollars in expenditure to develop spacecraft that were used once and then thrown away forever. NASA has clearly communicated they are no longer interested in “flags and footprints” style missions, but instead wish to foster a sustained presence beyond LEO. No matter what the specific plans are in store for the future, sustaining humans in space is unachievable without refueling in at least part of the picture.

“Reusability” of launch vehicles has already showed intriguing prospects in the commercial sector, and this trend is likely to continue for elements in space as well.

Identifying the most cost effective, feasible, and useful means of establishing a near-term propellant supply network is critical.

Whether the next destination for humans is a sustained lunar presence, near Earth asteroids (NEA) or Mars, the ability to routinely and affordably transfer propellants between a refueling module, just called a “tanker” here and thereafter, will significantly improve campaign prospects. With near-term propulsion technology, Mars transits with human-rated systems require propellant masses on the order of hundreds of metric tons[15]–[19]. A fully-fueled vehicle designed for such a mission cannot be launched by any single existing or near-term launch vehicle. Launch vehicle payload limitations necessitate the ability to efficiently launch, transfer, and store propellant for use in deep- space transit vehicles. Thus, distributing the propellant manifest over 10s of (relatively) inexpensive commercial launch vehicles (CLVs) is an attractive option.

New “heavy class” launch vehicles such as the Falcon Heavy and those in- development like the New Glenn and Vulcan Advanced Cryogenic Evolvable Stage

(ACES) present a new opportunity for providing a means of placing propellant in orbit

19 for use in transit vehicles. Of course, to get the propellant to its destination, a tanker is needed to do the following:

1. Integrate with a launch vehicle via a payload adapter fitting (PAF)

2. Store propellant in tanks until propellant is ready for transfer

3. Provide thermal management of propellant (heating for hypergolic propellant

and cooling for cryogens)

4. Separate from the launch vehicle and perform propulsive orbital maneuvers to

reach the required destination

5. Chase and either autonomously rendezvous and dock (AR&D) with the transit

vehicle, or establish a physical connection with the transit vehicle via some

other means without damaging either vehicle or causing disruptions in power

acquisition, avionics, communication, thermal management, life support, or

other critical systems

6. Settle tanks if necessary and establish a fluid connection with the transit

vehicle

7. Communicate necessary telemetry with mission control to confirm nominal

and safe operation

8. Once approved by mission control, open valves (and power on pumps if there

are any) and transfer propellant to the transit vehicle

9. Power down pumps (if applicable) and vent tanks if necessary

10. Separate and escape from the transit vehicle’s proximity

11. Loiter by performing orbital maintenance (if another fueling mission is

necessary)

12. Deorbit or transit to graveyard orbit after final mission termination

While satisfying all of these objectives, the tanker architecture must also be affordable, have acceptable risk, and meet the schedule demands of the mission or campaign. There are many ways to achieve all of these objectives, and some of the possible methods of doing so are analyzed in this document. Trades and sensitivity studies in tanker vehicle design and the fueling architecture were analyzed. The overall impact to several Mars human-exploration campaigns’ system mass, relative cost, robustness, complexity, and risk were all considered. The same or similar processes to the method used in this document could be applied to other exploration campaigns to Mars or other destinations.

1.1 Human Exploration Campaigns

1.1.1. Previous Mars Architectures

NASA’s Design Reference Architecture (DRA) 5.0 envisioned humans reaching

Mars either via nuclear thermal propulsion (NTP), traditional chemical propulsion, (CP) or a combination of chemical propulsion and aerocapture. The three-mission “long-stay” campaign included a crew of 6 with surface stays up to about one Earth year [15].

Dissimilarly, the NASA Jet Propulsion Laboratory (JPL) explored a “minimal Mars architecture” to reduce cost as much as possible. This architecture utilized hypergolic propellant for in-space propulsion (JPL also assumed the SLS Exploration Upper Stage

(EUS) could perform the Trans Mars Injection (TMI) Burn), and included a Phobos mission, a short surface stay, and a single year-long surface mission [19]. The Evolvable

Mars Campaign (EMC), like DRA 5.0, targeted long-stay missions, but utilized a hybrid

21 propulsion strategy: one vehicle with two separate propulsion systems; one capable of low thrust, high efficiency burns, while the other system was capable of (non-nuclear) chemical burns for high thrust, low transit time burns [17]. Typically, only low thrust burns are used for cargo delivery whereas high thrust is only used for crew transits. This hybrid vehicle uses both types of burns in tandem (a chemical burn for trans Mars injection (TMI), continuous low thrust during transit, and chemical Mars orbit insertion

(MOI)). For a summary of the various architectures, see Table 1 [15], [16], [20].

dual use use dual

950 950

- -

3/yr

SLS

Split

None

Surface

585 850

short stay) short

Unpres Rover Unpres

~130 t to LEO to t ~130

Batteries/Solar

Hypergols, EUS Hypergols,

4 (2 to surface for for surface to 4 (2 (crew) Propulsive

heat shield (cargo) shield heat

Phobos/Mars Short Short Phobos/Mars Long Surface/Mars

JPL Minimal Mars Minimal JPL

3 x 23 t for long stay long for t 23 3 x

1 x 23 t for short stay short for t 23 1 x

Aerocapture

800 1100

- -

km) 3/yr

SLS

Split

crew

surface

600

propellant

Propulsive

1000

3 x 18 t for for t 18 3 x

EMC Hybrid EMC

~130 t to LEO to t ~130

5 x 18 t for 1st 1st for t 18 5 x

SEP, Hypergols SEP,

subsequent crews subsequent

Pres Rovers (100+ (100+ Rovers Pres

Phobos/Mars Long Long Phobos/Mars

Ascent Vehicle O2 Vehicle Ascent

(1100 contingency) (1100

FSPS (40 kW total) kW (40 FSPS

dual dual

700 1100

- -

km) 3/yr

SLS

Split

Evolvable Mars Campaign Mars Evolvable

surface

400

propellant

SEP, ZBO ZBO SEP, LOX/CH4

1000

3 x 18 t for for t 18 3 x

EMC Split EMC

~130 t to LEO to t ~130

Aerocapture crews subsequent

Propulsive (crew) Propulsive

use HIAD (cargo) HIAD use

Pres Rovers (100+ (100+ Rovers Pres

Phobos/Mars Long Long Phobos/Mars

Ascent Vehicle O2 Vehicle Ascent

(1100 contingency) (1100

FSPS (40 kW total) kW (40 FSPS

5 x 18 t for 1st crew 1st for t 18 5 x

thrust

630 980

NEP

cargo)

Surface

Electric

Nuclear Nuclear

(crew and and (crew

Mars Long Long Mars

Low

Not addressed addressed Not Not addressed addressed Not addressed Not addressed Not addressed Not

dual dual

380 880

~12

km)

Split

crew

. Previously Proposed Mars Architecture Breakdown Architecture Mars Proposed Previously .

propellant,

S (40 kW total) kW (40 S

DRA 5.0 DRA

ZBO Cryogenic ZBO

2 x 40 t for each each for t 40 2 x

Aerocapture

Propulsive (crew) Propulsive

Crew Crew H2O/O2/N2

use shroud (cargo) shroud use (100+ Rovers Pres

Ascent Vehicle O2 Vehicle Ascent

Mars Long Surface Long Mars

FSP

Chem/Aerocapture

Table Table

dual dual

O2/N2

600

1000

9 6

km) 6/yr

rs Long Long rs

Split

crew

NTP

total)

(1000 (1000

170 t to LEO to t 170

Ares V Ares

Surface

400 ulsive (crew) ulsive

900

Cryogenic

propellant,

contingency)

ZBO Nuclear Nuclear ZBO

FSPS (40 kW kW (40 FSPS

2 x 40 t for each each for t 40 2 x

Aerocapture

Prop

Crew Crew H2O/

use shroud (cargo) shroud use (100+ Rovers Pres

~150 O2 Vehicle Ascent

Space Propulsion Space

anders

Duration in Free Space Free in Duration 1st for Launches Delivery Cargo Destination size Crew Duration Total Vehicle Launch Missi Surface Cadence Launch Max. In MOI L Power Surface Mobility ISRU

Figure 3. Explanation of currently planned SLS Block configurations.

Source: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170005323.pdf

Each Space Launch System (SLS) Block has a crew and cargo variant. Block 1 has the lowest performance of the variants, can only accommodate a single payload in a

5-m fairing, and incorporates the ICPS adapted from a Delta IV Heavy. Block 1B is similar to Block 1, but includes the EUS instead of the ICPS and an 8.4-m shroud, meaning a higher payload mass and volume accommodation. The crew version of Block

1B can also include a CPL. SLS Block 2 will integrate upgraded SRB motors for higher performance and a longer cargo fairing.

DRA 5.0 assumed an Ares V capable of delivering a ~150-170 t payload to low

Earth Orbit (LEO). The Space Launch System (SLS) has yet to fly, but it is expected that

SLS Block 1B and Block 2 will have much lower payload capacities than Ares V [21].

Recent cost projections and congressional testimonies by NASA also suggest SLS will

24 have a highly limited launch cadence in the future. According to NASA’s administrator, the current plan is to fly one SLS per year once the program is up and running [22].

Theoretically, this cadence could rise to two, possibly three launches annually, but no more is anticipated. Due to the specific assembly facility, number of launch pads, and crawlers available, SLS will never be capable of launch frequencies anywhere near as high as the commercial sector. Because many proposed campaigns over the years relied on now unrealistic SLS or Ares capabilities, these restrictions on launch vehicle performance and cadence necessitate significant architecture level changes.

In order to meet the gap forming between Mars campaign needs and SLS capabilities, [23] were one of the first to study augmenting DRA 5.0 with commercial and international launch vehicles in addition to SLS launches for the Mars transit vehicle

(MTV) launch and assembly campaign. Reliance on CLVs is a more attractive option for several reasons:

1. Cost: NASA would not have to invest in any additional launch vehicles or pay

for maintaining their continued operation. The only cost NASA would pay to

the CLV provider would be the “ticket to fly”. Prices for payloads on CLVs

are also considerably less expensive than government launch vehicles,

especially with the debut of new companies like SpaceX which are attempting

to close the recoverable and reusable launch vehicle systems business model

[24].

2. Schedule: CLV providers have already proven their high launch cadence

capability with SpaceX breaking its own record for most commercial launches

in a year (20 from 18). CLV cadence is only expected to improve with time.

3. Availability: Launch vehicles like the Falcon Heavy are presently available.

There is no need to design an architecture around the assumption that a bigger,

more capable rocket will be developed in time for the mission, when this may

never come to fruition. Launch vehicle development is a tremendously

expensive and lengthy process which typically ends behind schedule, over

budget, and below the required capabilities. Designing to existing launch

vehicles buys down risk significantly.

4. Reliability: CLV reliability has dramatically improved over the years. While

the reliability of SLS is expected to be sufficiently high, it is still more

uncertain than CLVs that have already flown multiple missions.

Incorporating CLVs for human exploration campaigns is slowly gaining headway.

NASA’s strategic goal # 3 calls for advancing commercial partnerships [25]. Given their high cadence and mid to heavy payload capabilities, CLVs are best suited for launching smaller elements like propellant tankers and logistics modules. Relying on CLVs for propellant tanker launches will be explored in more detail later.

1.1.2. Latest Mars Architecture

Similar to EMC, NASA is considering developing a Hybrid Propulsion System

(HPS) for transit to Mars [18], [26]. As previously explained, HPS relies on a combination of high and low thrust propulsion systems (SEP and chemical), thereby circumventing the need to develop multiple propulsion elements while still taking advantage of the high specific impulse of SEP. HPS nominally loiters in near-rectilinear halo orbit (NRHO) until it is loaded with enough propellant, crew, consumables, and logistics to journey to and from Mars. NASA is also planning to construct Gateway in

NRHO and send astronauts to explore the lunar surface in the near-term. Under current assumptions, Gateway and lunar surface exploration requires SEP and chemical propellant respectively (The power and propulsion element (PPE) on Gateway utilizes

Xenon hall-effect thrusters). Thus, for these notional concepts and programs, there exists a need to deliver SEP and chemical propellant to NRHO for Gateway, lunar, and Mars campaigns. This need can be met by a single tanker capable of delivering both of these propellant types in different ratios depending on the specific mission need.

1.1.3. Fueling Architectures

A fueling network would significantly improve and enable new missions. Fueling architectures have been studied time and time again, however, practically all proposed fueling architectures involve a depot [5]. Propellant depots are a powerful solution for space missions. Analogous to a gas station, the depot sits at some node in cislunar space and stores enough propellant to supply many missions. Also like a gas station, depots must be continuously operated, even if their only customer is using the asset once every

2.1 years to go to Mars. The concept simply does not meet the infrequent propellant demand. Although a depot would be wonderful to have, every element must also fit within the constraints of existing or near-term launch vehicles. Most proposed architectures, therefore, prescribe launching a monolithic depot on an SLS cargo variant to maximize the propellant mass delivered per launch. This investigation will consider an entirely different approach, for many reasons: SLS cargo flights have a highly limited cadence, and will likely be allocated for structures that simply cannot fit in a smaller fairing, like Mars transit vehicles, deep space habitats, large landers, etc. Additionally,

SLS cargo may not be a cost effective option compared to existing CLVs, especially

27 when propellant can readily be distributed over many smaller launches at a reduced price from government launch vehicles. Finally, even if you did launch one large depot on a cargo SLS, when all of its propellant is depleted, the depot will need a resupply of its own, which can be done with another costly SLS cargo launch, or with smaller tankers— which begs the question, why not design the network around small tankers in the first place, if they are what will inevitably be needed?

As previously stated, NASA entered the human exploration theater with the highly successful, but tremendously costly Apollo program. To this day, the culture of

Apollo still reverberates in the agency. NASA has recently announced a renewed interest in returning humans to the moon using an architecture with some major similarities to

Apollo. It can be tempting to use the Lunar Module (LM) as a rough template for designing a new architecture, because if it worked once it will surely work again, right?

Doing so is a logical fallacy for several reasons: 1. Apollo achieved mission success, but did so at great monetary cost. 2. NASA’s risk posture today is staunchly conservative compared to that era. 3. Apollo was designed as a sortie. All elements were discarded after a single, brief use for one mission. Refueling and reuse of Apollo modules was avoided, and subsequently, the Apollo architecture is highly incompatible with reusability.

Today, although NASA would like to procure easily reused and refueled elements for recurring missions, this will not happen without significant process and cultural changes. Instead of being an afterthought, the fueling architecture should play a pivotal role early on in the engineering design process, ideally pre-phase A. Such a change is not disruptive to the systems engineering process, but actually complements the process.

Trade studies and sensitivities are already a widely accepted part of systems engineering.

Instead of considering refueling as auxiliary to the system, it should be directly part of the system. Some awareness of a priori architecture impacts on refueling and vice versa should exist before down selecting architectures.

Previous fueling studies focused on depots, however there exists a new potential for less expensive options. Human exploration architecture impacts on commercially- enabled fueling architectures and vice versa have yet to be studied in detail. This investigation will target this gap, with emphasis on economically feasible concepts.

Refueling needs to be an important consideration in the decision analysis because it has one of the largest impacts on sustaining a campaign. Large campaigns inevitably require a physically achievable, cost effective solution for delivering tremendous propellant loads. Therefore, designing a successful fueling architecture is as much of a logistics problem as it is an engineering problem. The following section will investigate the logistics aspects of fueling as well as applicable logistics methods.

1.1.4. Logistics and Tracking

As already alluded to, fueling architectures are in many ways analogous to logistics networks omnipresent in modern society. Reliably supplying an extensive need, whether common goods to consumers or propellant to spacecraft, can be monitored effectively and improved by quality logistics practices. Perhaps the most common application of logistics is in the shipping industry. The cornerstone of success for companies like Amazon resides in their advanced shipping methods and sophisticated logistics capabilities. By harnessing automation and route optimization, millions of

29 products are shipped from factories to docks, properly loaded onto freighters and shipped across oceans, then redistributed in constantly morphing web until individual products reach the right customer. Countless man hours and funds are spent researching and designing logistics networks to improve reliability and efficiency. Numerous papers and patents all indicate a strong paradigm shift towards automated logistics [27]–[31].

The rise of CLVs and a need to use many of them for supplying large campaigns has created a new potential application for automated logistics. Generating satisfactory launch manifests and schedules for tens or hundreds of fueling and cargo flights can be significantly augmented with automation. Tracking manifested propellant and cargo throughout tens or hundreds of overlapping missions can also be automated. Automated manifesting and tracking is even more useful when the campaign or fueling architectures are still being traded. Sensitivities to the campaign can quickly be identified and cases can easily be tweaked if a new iteration is deemed necessary. With this flexibility, more time can be spent identifying the right solution rather than spending time crunching numbers manually.

An underlying assumption in a good logistics network is that nothing is perfect, and failures are an inevitability. The right solution does not constantly strive for the impossible goal of perfection, but instead assumes failures will happen, and when they do, the network can adapt to still succeed in the mission. Likewise, a good fueling architecture can adapt to mitigate risks—both programmatic and operational. A robust system has margin, and can overcome a few instances of failure so long as most deliveries are successful. [32] first investigated simulating the risks of orbital refueling and found concrete, feasible mitigations. Having at least one tanker as a contingency for a

30 delivery or launch failure significantly minimized mission risk. A few examples of programmatic and operation failures are shown in the table below.

Table 2. Potential Risks

Programmatic Operational  Tanker development schedule slips  Loss of tanker during launch

 Tanker dry mass requirements not met  Loss of tanker in transit

 Tanker is over budget  Loss of tanker before reuse

 Technology requirements not met  Failure to AR&D

(reliable AR&D, CFM, etc.)  Failure to fuel (within time required)

 Fluid transfer requirements not met  Failure to separate and escape from

vehicle

 Subsystem failure (comm., avionics,

CFM, power, etc.)

Significant research has gone into logistics supply for human exploration. Almost all of this area of research, however, has focused on supplying crew consumables and provisions and spare parts for environmental control and life support systems (ECLSS).

To the author’s knowledge, there is no presently existing, publicly available research which focused specifically on propellant supply from a logistics standpoint. As explained previously, there is a significant need for such an investigation. This is one of the major gaps in knowledge that will be addressed here going forward.

2. Methods

2.1. The Systems Engineering Process

Figure 4. The Systems Engineering Process Source: (Public domain) https://ocw.mit.edu/courses/aeronautics-and-astronautics/16-885j-aircraft-systems- engineering-fall-2005/readings/sefguide_01_01.pdf

The Systems Engineering Process is used across many industries to develop a product from what is needed. Working recursively, the systems engineering process is intended to bring high-level concepts to detailed and refined solutions. Each input, output, and segment of the cycle can be summarized in Figure 4.

2.1.1. Systems Analysis and Control

Control processes include risk management, scheduling, technical reviews, configuration management, and data management. Systems Analysis and Control is concerned with informing the decision making in systems engineering. Trade studies, feasibility analyses, and design analyses in the forms of modeling, simulation, and testing are all utilized in order to determine the impact of decisions on system requirements, risk, life cycle, and performance.

2.1.1.1. Trade Studies

Decision making is done in part with trade studies. Considering requirements, schedule, risk, cost, and lifecycle are the major considerations. Trade studies may: support functional analyses, define requirements, evaluate various architectures and concepts, inform material selection, examine design change impacts, and evaluate manufacturing processes, component standards, and system layouts.

2.1.1.2. Modeling and Simulation

Models are tools used to represent a system or process. Models can be mathematical, physical, or logical. The use of models to replicate a desired behavior is known as simulation. Modeling and simulation offer convenient solutions in the system engineering process while saving money, time, and improving the product.

Models can be divided into the following categories: virtual, physical, or a combination thereof. Physical simulations often provide the best results, but at higher cost. There is no substitute for physical simulation. However, virtual simulations offer a lower cost supplement that can be implemented earlier in the system engineering process.

The level of reliance on physical versus virtual simulation heavily depends on the required fidelity, the complexity of the system, the requirements of the system, the level of risk, and the system lifecycle. Typically for space applications, systems are highly complex and risk levels are stringent, meaning a heavy dependence on hardware testing is required. Hardware testing can also validate software simulations. When dealing with systems of systems (architectures) with high complexity, software is also an important part of systems analysis for making key high-level decisions when hardware may not be procured yet.

2.1.2. Systems Analysis in Practice

In practice, changes in requirements, capabilities, schedule, cost, and outputs from other loops being run in parallel to the systems analysis loop can all result in redefinition of inputs to trade studies. These new inputs, however, do not always arrive with ample time to generate the output before a decision has been made. This can cause decisions to be made a priori, which could impact the system performance in a detrimental way.

One mitigation is to streamline the systems analysis process in such a way that decision impacts can be predicted faster and with greater precision. Higher fidelity models and simulations can inform decision impacts with higher precision. Increased automation via programming scripts can provide ample trade study outputs much faster

34 than traditional “manual” methodologies. Faster iteration allows faster response to changes—both internal and external.

2.1.3. Architectures

Systems of systems (SoS), also referred to as architectures in the context of space missions is a relatively new discipline compared to traditional systems engineering.

Architectures cannot be examined and designed in the exact same manner as a single system. Architecture definition and refinement can be considered an auxiliary process completed in parallel to the systems engineering process. The output of each cycle can inform the inputs of the other cycle, or its own cycle [33], [34].

When designing space exploration architectures, one major consideration is the interdependencies of system designs on other systems. How do systems interface with each other? What constraints are present? How does system A affect the performance, cost, lifecycle, or risk of system B? Can System A be redesigned to better accommodate system B without a significant performance drop on system A?

The analysis approach to systems engineering becomes increasingly necessary for campaigns, or series of coordinated missions, involving scores of elements, each with their own systems and subsystems to consider.

Another major consideration is the technology implemented in the systems. An architecture can be designed around specific capability and vice versa. For instance, an architecture might rely on in situ resource utilization (ISRU) of some kind. Even a gravity assist can be considered an ISRU capability, and specific architectures may not be

35 able to complete the mission without the use of specific capabilities like a gravity assist.

Such capabilities are referred to as mission critical or mission enabling [33].

Inevitably, modifications to the requirements, lifecycle, performance, or cost of an architecture or capability will impact one another. Typically, architectures with low

Technology Readiness Level (TRL) introduce risk in the system because they have not yet been flight proven, and their performance and lifecycle are not fully understood.

Often, new space exploration and operations programs cannot be executed without at least some technology development. Technology investments must be made far in advance of phase B in a program so that the technology can be utilized on the mission.

Consequently, there exists high demand on systems analysis to predict which technology investments are feasible and worthwhile for future campaigns.

2.1.4. Architecture Definition

2.1.4.1. Space Technology, Capabilities, and Architectures

NASA plans to continue exploring destinations like the surfaces of the moon and

Mars with human astronauts—each campaign consisting of multiple related missions.

Each mission will require launch vehicles, transit vehicles, transit habitats, surface habitats, landers, ascent vehicles, and cargo to name a few. Technology investments must be made by the agency prior to the flight hardware development. Potential capabilities being considered for further development includes cryogenic fluid management (CFM), autonomous rendezvous and docking (AR&D), in-space refueling (iSR), in-space servicing and assembly (iSSA), and electric propulsion (EP) among many others.

NASA uses a space technology roadmap (STR) to trace key plans for investment with desired future missions demonstrating the technology [1]. The choice in architecture impacts the technology development roadmap and vice versa.

Strategic investments in space technology are ideally posed years or decades in advance. To be an effective use of resources, typically only mission-critical and mission enabling technology is developed beyond TRL 1-3. Technology investments, as well as missions themselves, may change over the years. Architecture trade spaces are readily influenced by projected capabilities. Architecture trade studies can also reveal high-level technology and capability requirements, which provide insight for further studies and future capability investments.

2.1.4.2. Architecture Trade Studies

The goal of architecture trade studies is to identify an architecture that relies on feasible technology, and meets the cost, schedule, and risk requirements. Architecture trade studies can be used to generate potential options and identify which architectures offer the best balance of needs. Any number of architectures may be considered before a final option is selected. For campaigns involving several missions, each with a multitude of options for how the target destination is reached, what will be accomplished once at the destination, and for human-rated missions, how long will they stay, and how will they return to Earth, the number of architecture options can become unbounded.

In the past, the scope of architectures considered had to be limited to what could be developed ‘manually’ with the workforce and time available. Thanks to the advancements made in computing power, it is now possible to develop scripts that

37 generate virtually infinite architecture options in a fraction of a second. By using this amount of information properly, more favorable design choices can be made.

Despite a virtually infinite number of potential options, there is limited time to explore usually a handful of architectures before a decision has to be made. Mission design must be optimized for each option, elements should be sized, and estimates on cost, performance, risk, and expected lifecycle must be made.

Pre-phase A trade studies are not the time for detailed engineering design down to the subsystem or component level. Instead, physical models and many approximations based on previously flown missions are utilized to create parametric models of elements.

The models are typically virtual, comprising of system mass, cost, risk, and performance estimates, which are used to compare architecture options against each other.

The growing complexity of human exploration campaign architectures warrant the development of virtual simulations to represent and compare many potential architecture options in minimal time. By developing software to increase trade study automation, more architectures can be developed than was previously possible by hand. It could be possible to generate millions of architecture options and impose constraints in order to narrow down the options to an “optimal” architecture.

The entirety of this investigation resides within pre-phase A of the systems engineering process: concept studies. Typically, mission concepts are identified, their performance and feasibility is evaluated using models and simulation.

1.1.4.3 Spacecraft Sizing

Mass estimation and forecasting has been a part of spacecraft design since at least the days of Mercury Program [35]. The concept of sizing relies heavily on both physical models and tabulated databases of heritage systems. Spacecraft design is fundamentally an iterative process, thus mass estimates much first be used as placeholders to represent systems and subsystems until they are designed in detail. With these estimates, high-level design decisions can be made such as down-selecting architecture options. The design and sizing process is summarized in the table below1:

Table 3. Summary of the Spacecraft Design and Sizing Process 1. Define basic requirements and constraints

2. Using 1., select preliminary architecture and high-level configuration

3. Establish delta-v, mass, and power budgets

4. Design preliminary subsystems

5. Design baseline spacecraft configuration

6. Using output obtained from the above steps, return to 1. and iterate again with

updated inputs. Requirements, constraints, and budgets can and will always

change.

This is a highly generic and simplified version of the process, but the general process of using this iteration loop works for any missions and campaigns.

1 Table 3. was generated using [36] as a reference. 39

2.2. Propellant Tracking Theory

2.2.1. Capabilities and Rationale for Propellant Tracking

Campaign architectures are not simply designed once. The process is recursive.

Each architecture output is studied in detail in order to identify what changes should be made for the next iteration, and what conclusions can be made about capabilities that need to be developed to complete the missions. One way an architecture might be studied is a form of simulation known as tracking. Once elements are modeled and a mission designed, it is necessary to track where each element is physically for the duration of the campaign. Other useful information includes representations of every element’s expected structural and thermal loads, power, propellant, and cargo loads, interfacing with other elements, and so on and so forth for the entire campaign.

Having a record of this information helps answer useful questions about the architecture. Again considering a campaign consisting of tens or even thousands of elements, where many of them may interface with each other for fueling, repair, or crew transfer, all elements must be tracked to ensure, for instance, two vehicles will be in each other’s vicinity at the right time for a rendezvous. Another example might include using tracking to ensure a propellant tanker will be capable of transferring enough propellant in time for an element to complete the required mission even if, for instance, a docking port in close proximity to the tanker is being used by a third element at the same time. Thanks to modern computing, these constraints are easy to impose against virtually any number of elements in the campaign.

The scripts developed and discussed in future sections tackle the specific task of propellant tracking. Although tracking logistics, power levels, thermal loads, and many other parameters could be accomplished utilizing modifications to theses scripts

(discussed in more detail later), automated propellant tracking was identified as the most essential and urgent need for several reasons: the latest Mars campaign architectures being traded have all required fueling of transit elements using propellant tankers. The tankers, often the most numerous of elements in the campaigns, significantly impact cost.

Another dimension of complexity was also added by the prospect of utilizing a hybrid propulsion system (HPS), which would burn two types of propellant: one chemical propellant and one solar electric propellant (SEP). Both propellants would need to be tracked simultaneously.

An architecture utilizing hybrid propulsion would add robustness to transit vehicles and forgo the need to develop two separate propulsion systems for crew and cargo transport. For the same reason NASA is considering HPS, tankers would also be capable of carrying two types of propellant. With slightly oversized tanks, the ratio of

SEP to chemical propellant can be adjusted for each launch to meet the immediate propellant need throughout any point in the campaign. Tankers with significant fuel leftover could be used to fuel HPS again before it departs on another mission.

2.2.2. Program Logic and Structure

After each fueling event, the fuel load in both the tanker and the propulsion element has changed, and must be recorded. This, in turn, changes the current propellant need, prompting another decision to be made: to what levels should each propellant be loaded onto the next tanker? The object of the script is to automatically decide the correct

41 propellant levels that should be launched on a tanker to minimize the total number of tankers necessary to fuel a campaign. Once a fueling architecture has been generated, the script also produces graphical representations of the propellant being tracked between tankers and propulsion elements. A breakdown of the decision logic is shown in Figure 5.

Figure 5. PropTracker Basic Code Structure

One example of a visual depiction of propellant tracking is shown in Figure 6.

Clearly, calculating the propellant loads for each of ~50 tankers was tedious and time consuming. Additionally, with manual calculation, little time was available to understand campaign impacts on and sensitivities of the fueling architecture. In fact, just in the timeframe of this investigation, multiple launch vehicle payload projections, lander mass,

HPS tank size, and many other parameters of the campaign changed. Knowing the impact

42 of these changes by clicking a button is far more advantageous than running the propellant load calculations manually.

PropTracker was developed to track two quantities simultaneously. Those quantities were originally intended to represent:

1. The mass of a solar electric propulsion (SEP) propellant in any given element

2. The mass of a chemical propellant such as a cryogen (e.g. hydrogen) or

hypergolic propellant (e.g. hydrazine) in any given element

This would be useful for campaigns involving two propulsion systems, a high thrust chemical propellant for fast transfers with crews, and high efficiency SEP for cargo transfers.

PropTracker can also be useful if tracking or manifesting is not only constrained to one SEP and one chemical propellant. The variables for mass tracked could represent any type of fuel or oxidizer or even logistics (as will be discussed in later sections). In this sense, PropTracker was designed to offer flexibility to the user.

Figure 6. PropTracker Sample of Graphical Output

2.2.3. Program Verification

PropTracker simulates a simple, repetitive process. A propellant need is given, and constraints on the delivery options are imposed. The decision logic dictates the propellant load to meet the need with as few launch vehicles as possible. The propellant is loaded onto a tanker, what is needed is transferred to an element, then the loads for tanker and element are stored in the program’s memory. From there the process repeats until the campaign propellant loads are all computed and tracked. The data verification of

PropTracker is commensurate in simplicity. A few key checks based on physical constraints are needed for each case run using the tool:

1. The SEP and chemical propellant loads cannot exceed the maximum space

physically allocated in their tanks.

2. The gross mass of the tanker cannot exceed the available payload capacity of

the launch vehicle selected for tanker delivery.

3. The tanker must either:

a. Transfer all of its propellant to the element (minus ullage)

b. If the tanker load exceeds the propellant need, transfer only the

propellant need (remainder is transferred to the next element)

4. Propellant load of each element must equal its respective propellant need.

To verify PropTracker output, simple checks on these above constraints were performed with an automated method. When each case is run through PropTracker, a report is generated which contains propellant needs and loads as well as models of the launch vehicles, tankers, and elements. This report is printed to an Excel workbook specifically designed to reconstruct the fueling events line by line. Each row has logical check for each of the four constraints above. If a constraint violation is found for any case, the workbook produces a “red flag”, which is made obviously visible to the user.

The workbook is designed so that if one event triggers a red flag, all preceding events produce red flags. This verification ensures each simulation satisfies the four most basic constraints (listed above), and facilitates quicker debugging when new features are added to PropTracker. The report also serves other important purposes. The workbook can be saved in case the same analysis must be repeated in the future, or the user wants to go

45 back and see the full details of an archived case. Recording the version number of

PropTracker in the report is also essential for analysis documentation purposes. If an analysis is completed, new features are added, then a bug is found, one must verify the cause of the bug and be able to verify the bug did not affect the previous analysis. These fundamental “best practices” of programming and simulation are essential for verifying data.

2.3. Campaign Analysis

Before element requirements are ever set, their intended mission must be clearly defined. The mission usually starts with a simple goal, e.g. “bring humans to the surface of the moon and back to Earth.” Then, as discussed in Section 2.1.1, the systems engineering process involves systems analysis. Trade studies and performance evaluation are two pieces of systems analysis that shape the system architecture.

In this context, the system architecture is the campaign of missions being considered. The trade studies will involve a “baseline” architecture, and several “deltas” or perturbations to this system architecture. The performance of the campaign will be assessed in the results and discussion section. Lunar and Mars campaigns will be considered. Deltas will include payload requirements, launch vehicle fleets utilized, propellant choices, CFM options, tanker design, and propellant transfer rates. The campaigns themselves were used as inputs to the analyses and reproduced below with permission. Each of the following sections will summarize the assumptions, the baseline architecture, and the trades considered. The results of each trade will be in Section 3.

2.3.1. Mars Hybrid Propulsion System Campaign

Table 4. Hybrid Propulsion System Mass Breakdown

Category Mass (kg) Structure 5,805 Thermal Control 2,036 Electric Propulsion 4,116 Chemical Propulsion 3,010 Electrical Power 6,299 Avionics & Control 138 Growth 6,426 Dry Mass Subtotal 27,830

Max Xenon Load 34,000 Max LOx/LCH4 Load 24,000 Max Wet Mass 85,830 Deep Space Habitat Payload 45,000 HIAD Lander Payload 53,000 Figure 7. Hybrid Propulsion System

Source:https://www.nasa.gov/sites/default/files/th umbnails/image/deep_space_transport.jpg The HPS is one potential option for a Mars in-space transportation system.

Instead of developing two separate propulsion systems for slow and fast transfers respectively, a single element capable of both types of propulsive maneuvers could be utilized instead. HPS (and each tanker) would require some systems unique to each type of propellant, such as two propellant storage tank types: one designed for SEP and one for chemical propellant. The HPS campaign was not designed by the author, but used as an input for fueling trade studies in this text. A summary of the HPS model used for this investigation, presented in [37], is reproduced in Tables 4, 5. The details on the design of the HPS campaign can be found in [18], [26], [37]–[39].

Table 5. Baseline HPS Summary

Features Design Life 5,500 days Element Diameter 5.62 m Element Length 9.9 m Main Propellant Type Xenon O2/CH4 No. Engines; Engine Type 8 x 50 kW Hall 6 x Cryogenic Thrusters Engine Thrust (100%) 4.5 kN Each Engine Isp (100%) 2,600 s (800 V) 351 s No. of Engine Restarts 10+ 15+ No. of Tanks 12 x Xenon 1 x O2; 1 x CH4 Tank Material COPV Al / Ti RCS Propellant

No. Engines; Engine Type 52 x Supercritical O2/CH4 Engine Thrust (100%) 445 N Each Engine Isp (100%) 330 s Power System Solar Arrays 300 V Rollout + 120 V Body Mounted BOL Generation 675 kW Main + 1.4 kW Commissioning Structure ISS SARJ Gimbals Cell Type; Efficiency IMM Solar Cells; 32%

Because of launch vehicle payload limitations, HPS would need to be launched only partially fueled. Tankers could then fuel HPS with different ratios of SEP or chemical propellant depending on the specific mission opportunity and payload. Another key aspect of HPS is the system reusability. Each HPS element is designed with enough propellant to return to cis-lunar space “empty” so it can be refueled and reused for the next mission in the campaign. Even the cargo delivery elements are returned to considerably reduce the number of HPS elements that need to be built and thus reduce the number of SLS cargo launches needed as well.

Two HPS designs will be considered: a Xenon-hypergolic version and a Xenon- liquid oxygen, liquid methane version. The two HPS versions were developed years apart from one another and cannot be compared to each other because each version was developed under different assumptions. However, different sensitivities and trades were conducted in parallel for each HPS design. A summary of the baseline Xenon-liquid oxygen, liquid methane version is shown in Tables 4, 5. The hypergolic version will be discussed in more detail in future sections.

The following are the assumptions for this system architecture:

1. A 9:2 lunar synodic resonance2 Near Rectilinear Halo Orbit (NRHO) is the

staging point for all fuel transfers and loitering of elements3.

2. Once HPS returns from Mars, it performs an NRHO insertion burn and loiters

until it is refueled and reused for the next mission.

3. Several identical HPS elements must be launched in order to deliver all crew

and cargo for each mission.

4. A lifetime of 15 years was assumed for a given HPS. No iSSA was assumed

to build up or maintain HPS elements beyond propellant and logistics

resupply.

5. Each HPS launches partially fueled on an SLS Block 2.

6. Each tanker is launched via commercial launch vehicles (CLV) on a Ballistic

Lunar Transfer (BLT) requiring no insertion burn (0 m/s) after the trans lunar

2 9 revolutions per 2 lunar months on average (or 6.562 days per revolution) 3 Note: “HPS element” refers to a single hybrid propulsion system element. Each campaign has many identical HPS elements. “Element” may refer to any vehicle or distinct system not permanently integrated with another system or vehicle or module. E.g. tanker, HPS element, or a deep space habitat. 49

injection (TLI) burn performed by the upper stage of the CLV (However,

tanker RCS tanks were sized for a 5 m/s course correction burn and a small 10

m/s insertion burn as a contingency).

7. Each cryogenic version of HPS is assumed to be fully capable of zero boil-off

(ZBO) active CFM.

8. Unless otherwise specified, each tanker is assumed to be capable of active

CFM (ZBO). A cumulative 3% of each tanker’s propellant load was assumed

to be lost before transfer due to boil-off during launch and ullage from

transferring propellant.4

9. HPS:

a. 675 kW beginning of life (BOL) solar arrays

b. 400 kW SEP thruster power

10. Mid-latitude (19 deg.) target for all landers

11. If after fueling an HPS element a tanker still has enough leftover propellant to

transfer (>10 kg of either propellant), the tanker will loiter in NRHO until it

provided a second propellant transfer to another HPS element.

12. Mission timeline in the mid-2030s

a. 2031: Deep Space Transit (DST) Habitat deployed and crew checkout

b. 2033: One-year shakedown mission in HEO

c. 2035: Depart for three-year orbital mission

d. 2039, 2043, and 2048: Depart for Mars surface mission

13. Crew of four

4 This percentage of propellant lost was chosen as a starting point, not a precise estimate. Determining the precise “loss factor” was not a major focus of the investigation, but could be future work. 50

14. Conjunction class missions for surface missions

a. Minimum 300 days in Mars vicinity

b. Maximum 1,100 days round trip missions

15. Crew and cargo reach Mars separately

16. SLS Block 2 Crew: Orion plus 9 t co-manifested payload

17. SLS Block 2 Cargo: 45 t TLI (C3 = -1)

18. Payload Adapter Fitting (PAF) not included in payload capacities

a. 2.5 % mass per payload manifested on an SLS

b. 1 % mass per payload manifested on a CLV

19. Minimum of 180 days between Mars missions for checkout and aggregation

20. Once HPS delivers a lander to Mars orbit, it returns to NRHO “empty” for

refueling and reuse

21. Cargo and crew trajectories are similar, with the latter being slightly faster

thus requiring a greater ratio of chemical propellant to SEP

22. As solar arrays degrade (assumed 1.5% loss in power per year in space), SEP

engine performance declines and more chemical propellant is needed to close

trajectories. Near the element end of life, closing trajectories becomes

increasingly difficult and simply not possible (especially for heavier payloads)

depending on the opportunity year and specific trajectory (crew or cargo).

The propellant needs were based on the campaign shown in Figure 8. HPS was sized parametrically with an in-house sizing tool based off physical models and data from proven flight systems and subsystems. The design has been refined over the years of

51 work put into the system. All crew and cargo trajectories were optimized in Copernicus.

The capabilities of the CLV options assumed are shown in Table 6.

2031 Shakedown 2035 Mars Orbital 2039 Mars Surface 1 2043 Mars Surface 2 2048 Mars Surface 3

Figure 8. Mars Campaign to be used as a basis of comparison

Table 6. Launch Vehicle Capability Assumptions5

TLI Performance Full Usable Name Provider Payload (t) Payload (t) SLS Cargo NASA 45 43.9 Falcon Heavy Commercial 16.8 16.6 New Glenn Commercial 14.9 14.8 Vulcan ACES Commercial 14 13.9 Ariane 6 Commercial 10.6 10.5 Vulcan Centaur Commercial 9.5 9.4

5 Payload performance of CLVs was compiled based off publicly available data. Performance to orbits not available publicly was interpolated based on data which was available. Note these assumptions apply to the Mars campaign only. Slightly different CLV performance estimates were used for the lunar campaign because different information was available at the different times each of the studies were conducted. 52

NRHO offers low orbital maintenance (~10 m/s per year) and can be reached via ballistic transfers to eliminate the need for an insertion burn. NRHO is an L2 southern halo orbit in the Earth-Moon Lagrange system. For visualization purposes, see the proposed NRHO orbit for Gateway in Figure 96.

Figure 9. Visualization of NRHO proposed for Gateway

Source: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20190030294.pdf

NRHO can be reached via traditional fast transfers or ballistic lunar transfers

(BLTs). The latter targets the edge of the Earth’s sphere of influence and the ballistic payload then uses the sun to perturb its orbit in order to achieve insertion without an insertion burn. The disadvantage to BLTs is they are much slower than fast transfers, typically over 100 days compared to a ~4 day fast transfer. BLTs would not typically be useful for transferring crew, but unmanned elements, cargo, and propellant can all be sent

6 In the figure, the orbit appears to wobble each period because of the ephemeris model used. A detailed explanation of this behavior is provided in [40].

53 in this fashion. In the case of slow transfers of propellant, active CFM is needed for cryogens, otherwise all the propellant will boiloff in transit.

For the trades considered, the fueling architecture was kept constant. Other campaign variables were changed, and the impact to the fueling architecture was observed. The tankers are delivered by CLVs to TLI. The tankers then reach NRHO via a

BLT. Once they arrive at NRHO, tankers perform AR&D with the HPS elements then begin fueling. Next, the tankers escape and deorbit if they have no propellant remaining.

If they do have left over propellant, they can loiter in NRHO, using ZBO cooling to maintain their propellant until anther HPS element is launched or returns from Mars. The tanker design is summarized below:

Table 7. Passive and Active TCS Hybrid Tanker

Subsystem Breakdown Mass with Growth (kg) Structures 2,409 Propulsion 250 Power 411 Avionics 567 Thermal 791 Fluid Transfer System 444 Dry Mass 4,872 Non-Propellant Fluids 44 Bus Propellant 1,035 Total Stage Gross Mass 5,950 Maximum Supplied Fluids 12,784

2.3.1.1. Campaign Trades

The trades studies in this section will be different from the lunar campaign. The goal here will be to keep the Mars transit architecture static while varying the fueling architecture to see the impact on the entire campaign. The following trades were made for the HPS Mars campaign (to be explained in more detail in their respective sections):

1. Lander options: hypersonic inflatable aerodynamic decelerator (HIAD) landers

were assumed for the baseline campaign. A delta to the baseline was assessed by

instead using mid-range lift to drag ratio (Mid-L/D) landers, which can deliver

heavier payloads, requiring more propellant for the campaign.

2. Commercial launch vehicle fleets: Many commercial launch vehicles were

considered as options for launching tankers in the campaign.

3. Active versus passive CFM: The cryogenic LOX-methane HPS was used to

conduct a trade on CFM technology in propellant tankers.

4. Propellant transfer rates: A first-order estimate of propellant transfer rate

requirements for various campaign options was derived based on scheduling

needs.

5. Tanker design: The baseline tanker was compared to a parametrically sized

reusable “tanker bus” option which delivers minimalistic tankers.

2.3.1.2. Hypergolic Chemical Propellant Alternative

DRA 5.0 Addendum 2 featured a Mars transit vehicle similar to the current HPS, but instead relied on hypergolic propellant and SEP (Xenon) propellant (instead of cryogenic and SEP propellant). For clarity, this version of HPS will be referred to as the

“storable HPS” in this text. The assumptions made in this study were significantly different from those made for the more recent liquid oxygen-methane HPS previously

Table 8. Mars Crew Campaign: Storable HPS

Mission Opportunity Mission Type

2033 Orbital

2037 Surface Long Stay

2041 Surface Long Stay

2045 Surface Long Stay

Table 9. . Storable HPS Mass Breakdown Category Mass (kg) Structure 6,300 Thermal Control 1,270 Electric Propulsion 4,100 Chemical Propulsion 1,700 Electrical Power 6,640 Avionics & Control 180 Growth 2,660 Dry Mass Subtotal 22,850 Max Xenon Load 23,400 Max LOx/LCH4 Load 18,600

Max Wet Mass (Excluding Pyl.) 64,850 Deep Space Habitat Payload 48,000 Figure 10. Storable HPS Model. Mars Lander Payload 46,000 Source: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170009 117.pdf

discussed. Because each architecture had different assumptions, it would be illogical to compare the two campaigns. However, the hypergolic-SEP architecture is still feasible, and fueling related trades and sensitivities can still be a useful consideration.

In order to understand the sensitivity of the storable HPS vehicle design on the fueling architecture, a campaign was first defined (see Table 8). The parametric model of the storable HPS presented in [38] and used for this analysis is summarized in Tables 9,

10.

Table 10. Storable HPS Key Information

Features Design Life 5,500 days Element Diameter 6.2 m Element Length 7.4 m

Main Propellant Type Xenon N2O4/MMH No. Engines; Engine Type 24 x 13.3 kW Hall 10 x Aerojet R42DM Thrusters Engine Thrust (100%) 890 N Each

Engine Isp (100%) 2,000-3,000 s 327 s No. of Engine Restarts 10+ 15+ Tank Material COPV Al / Ti RCS Propellant

No. Engines; Engine Type 32 x Astrium S22-02 Engine Thrust (100%) 22 N Each Engine Isp (100%) 285 s Power System Solar Arrays 300 V MegaROSA + 120 V Body Mounted BOL Generation 500 kW Main + 8 kW Commissioning Structure ISS SARJ Gimbals Cell Type; Efficiency Li-ion, 23.8kWH at 28 V

2.3.1.3. Power Trade Space

The limiting SEP performance parameter is the power available. When HPS reaches Mars, the added distance from the sun reduces the SEP performance, requiring the CP system to compensate. The CP system must also bear a greater burden as the arrays degrade and power available is reduced over time. To understand the full trade space, several parameters were varied: the solar array output power was varied to see how changes in power available would affect propellant required, and thus the number of refueling flights required. The SEP power was also varied to study the same effects. This sensitivity assumed the dry mass did not change; in other words, the array efficiency increased or decreased, not the size of the arrays themselves. The expectation is that

57 greater power would allow the simulation to rely more heavily on the SEP system, so the

SEP propellant need would rise and the CP propellant need would fall. Using the campaign assumed (crew flights only, no cargo), the fueling architecture sensitivity was analyzed by varying the following:

 Array beginning of life (BOL) power range: 400 - 600 kW

 EP power: 265.3 – 504.07 kW

 Tanker payload propellant mass fraction

Because obtaining propellant needs requires running trajectories for each of the many cases, only crew trajectories could be considered with the time available. A supposed “crew-only campaign” was constructed. Although not entirely representative of a campaign, the analysis is still justifiable for two reasons: 1. The comparisons made with the campaigns are all relative. As long as they have the same, reasonable assumptions, the desired deltas can be studied. 2. The driving cases are the crew trajectories because they are more constrained by time of flight. Cargo trajectories could be studied as well, but their sensitivity is less critical to understand.

2.3.1.4. Latitude Sensitivity

The baseline assumed delivery of all landers and surface assets to a single, mid- latitude location. With the current habitat mass and lander masses, the campaign could theoretically target a single location across the entire campaign located at higher latitudes. Trajectories were computed to obtain propellant needs across a range of latitudes, and the impact of destination latitude on fueling flights was investigated.

Destination latitude does not impact the inbound or outbound trajectories, but an apotwist

58 maneuver must be done in Mars orbit in order to target the desired latitude. The delta-v varies, but generally increases for higher latitudes. This impacts the propellant needs of the campaign.

2.3.1.5. Electric Propellant Options

Xenon, only making up 0.000009 % of Earth’s atmosphere, is of limited abundance, has high demand across numerous industries, and requires high amounts of energy to process [41]–[43]. A campaign requiring approximately 100,000 kg of Xenon for each crew transit (like the HPS and NEP campaigns evaluated later) will indubitably be expensive no matter what. However, using this sheer quantity of Xenon may significantly drive the campaigns’ affordability. Fortunately, there are alternatives, however. Krypton, another noble gas, has similar performance to Xenon at a fraction of the cost of Xenon. Xenon is currently preferred to Krypton in part because for low power thrusters (~1 kW), the discharge efficiency gap is significant. However, Laboratory experiments comparing performance of the NASA-457M 50 kW Hall-effect thruster

(HET) showed negligible changes in thruster efficiency between Krypton and Xenon at high voltages [44]. Krypton also has theoretically ~20% higher specific impulse than

Xenon, and comparable thrust can be achieved at similar operating conditions to Xenon based on the experimental results of [42]–[44]. Krypton also has similar ionization energy, meaning few, if any, propulsion system modifications would be necessary.

(Bismuth, on the other hand, has even lower ionization energy, but the metal must be vaporized for ionization. This extra step requires an extra subsystem and more energy, whereas Xenon and Krypton gas feed systems are lower complexity. Not to mention, a vaporized propellant may cause additional deposition problems on the space vehicle not

59 already seen by the noble gas propellants [42]). Thus, Krypton may be a viable alternative worth examining for future architecture level performance analysis.

2.3.1.6. Tanker Design Sensitivity

The baseline campaign was evaluated across a range of tanker payload propellant mass fractions (PMF). For 50 different payload PMFs ranging from 0.5-1.0, the fueling was calculated with PropTracker for several different CLV options. Because the tanker model in this case was just a mass fraction, no maximum packaging limit was imposed on either propellant based on tank volume. The only loading constraint imposed for these cases was the tanker gross mass could not exceed the CLV payload capability. All other assumptions were the same as those described in previous HPS campaign trades.

2.3.1.7. Lander Payload Growth Sensitivity

Surface assets and crew can be delivered to the surface via several lander designs.

One of which is a HIAD lander, which is designed to save mass and volume via an inflatable aerodynamic decelerator. The HIAD was the lander design assumed in the baseline. Alternatively, a Mid-L/D lander could instead deliver assets and crew to the surface. This monolithic lander has a higher mass and volume, but could offer advantages over the HIAD lander, making it worth conducting a trade study.

For this trade, the entire baseline campaign was assumed, except all landers in the campaign were replaced with a Mid-L/D design. The Mid L/D gross mass assumed was

67,000 kg whereas the baseline HIAD lander gross mass was 53,000 kg [45]. To deliver the Mid L/D lander to the NRHO staging orbit, SLS Block 2 throws the lander to an altitude of 13,000 km, then two CLV upper stages “boost” the lander on a TLI. (The

HIAD lander only required one CLV boost stage). Maintaining the same SLS launch cadence assumed in the baseline campaign was a priority. From the heavier lander mass, the entire campaign was reassessed to determine the new trajectories, obtain propellant requirements from the new trajectories, and determine updated departure dates, etc. Once the propellant needs and departure dates were determined, fueling scenarios were obtained.

2.3.1.8. Commercial Launch Vehicle Trade

Several heavy lift commercially available launch vehicles are available or are expected to be available options in time for a Mars campaign beginning in the mid to late

2030s. Because each CLV’s availability and launch cadence is uncertain, it is extremely beneficial to know how these variables affect a campaign that partially or fully relies on

CLVs. Table 6 summarizes the estimated payloads of various CLVs to TLI. These estimates serve as the assumed performance for this trade and every trade unless otherwise specified.

Because the payload to TLI of each CLV is an estimate, they should not be taken as absolute and necessarily accurate values. CLVs still in development have even more uncertainty with their payload capacities, which should be taken into consideration when interpreting this analysis.

2.3.1.9. Tanker Thermal Control System Trade

The baseline fueling architecture assumed tankers, equipped with active ZBO thermal control systems (TCS), would be launched to TLI on a ballistic transfer. This transit can take roughly 120 days (depending on epoch), but the advantage is the insertion

61 burn into NRHO is typically negligible (<10 m/s and depending on epoch can be larger).

This allows tankers to be designed with no main propulsion system (MPS) and only a reaction control system (RCS). This saves mass and increases the amount of propellant that can be delivered by each tanker. However, the TCS required to achieve ZBO for several tons of cryogens has a high degree of uncertainty in the mass and power requirements, simply because such a system on this scale has to fly. The more power required to prevent boiloff would mean the larger the tanker’s solar arrays, thus driving up the inert mass and reducing transferrable propellant. Of course, there is a simple alternative that would circumvent this problem altogether: a fast Hohmann transfer (4-5 days) to NRHO. Assuming fueling can take place shortly after NRHO insertion, a passive

TCS consisting of multi-layer insulation (MLI) and other systems requiring no power can reduce boiloff for the short time of flight until fueling occurs. The downside, however, is that fast transfers require a significant insertion burn, meaning some of the tanker mass must be dedicated to a system, either MPS or RCS plus propellant mass required for this burn. Another significant source of uncertainty exists in the level of boiloff that can be achieved with passive methods for the mission duration. Somewhere, a “break-even” point exists between the slow transit architecture (actively cooled tanker with ZBO) and the fast transit architecture (passive tanker TCS with boiloff).

Two fueling architectures were compared with sensitivity considerations for each: the propellant mass fraction (PMF) of the slow transit tanker was varied while the fast transit tanker boiloff rate was varied. The goal was to determine what active TCS mass efficiency would be required in the slow transit architecture for a given boiloff rate in the fast transit architecture to make a ZBO tanker justifiable.

The same baseline commercial hybrid tanker (ZBO) was used for the slow transit case. The fast transit tanker was sized by adjusting the dry mass of the baseline tanker.

Cryocoolers were removed, array size was reduced, and all other systems were kept the same. The system mass comparison can be seen in Table 11. The following additional assumptions were made:

1. The propellant needs were taken directly from the baseline HPS campaign

detailed in previous sections (medium HIAD landers delivered to the mid

latitudes in the 2030s)

2. Slow tanker: In addition to the unusable propellant (residuals and ullage), it was

assumed that 3% of all usable propellant could not be transferred due to losses or

inefficiencies in the propellant transfer feed lines

Table 11. Passive CFM Hybrid Tanker

Mass with Growth (kg) Active TCS Passive TCS Subsystem Breakdown Tanker Tanker Structures 2,409 2,284 Propulsion 250 250 Power 411 218 Avionics 567 567 Thermal 791 362 Fluid Transfer System 444 444 Dry Mass 4,872 4,125 Non-Propellant Fluids 44 44 Bus Propellant 1,035 2,573 Total Stage Gross Mass 5,950 5,516 Maximum Supplied Fluids 12,784 12,784

2.3.1.10. Propellant Transfer Rate Trade

A notional fueling concept of operations and notional launch schedule provided a basis for estimating required propellant transfer rate. It should be noted that this is not a study intended to set requirements themselves given the conceptual, first-order nature of

63 the study. Instead, determining a rough estimate of the propellant transfer rate required for this campaign can help in making an architecture selection. At the very least, this high-level analysis should first verify the transfer rate is reasonable with current or near- term technology. If this is the case, it is also useful to estimate whether transfer rate is a driving performance metric. Essentially, what this investigation will identify is: 1. Is the required transfer rate feasible? 2. What is the likelihood of not meeting this requirement?

A quick assessment of the results should roughly approximate the level of likelihood of this risk.

The assumption for this assessment was that each tanker fuels the elements one at a time, transferring one type of propellant at a time, no simultaneously. From the trajectory optimization, the arrival and departure times for all of the elements was known, so the window of time where each element is in Earth’s vicinity before traveling on the next mission completely dictates the transfer rate. Additionally, a time penalty of 18 hours per tanker was assumed for proximity ops, autonomous rendezvous and dock, checkouts before initiating fluid transfer, undocking, and escape time.

2.3.2. Lunar Exploration Campaign

PropTracker can be used to assess the impact of vehicle and architecture aspects on the fueling campaign. In this section, the fueling tanker vehicle design and architecture will remain static while a lunar surface Human Landing System (HLS) is varied. Then, HLS will be kept static and the fueling architecture will be varied. The goal is to identify which HLS architectures and which fueling architectures are better suited for reuse.

Many feasible fueling options exist for HLS, but there are two main schools of thought that affect the architecture dramatically: the level of reusability, and the origin of the propellant itself. Whether the propellant is derived from Earth or excavated and processed on the lunar surface heavily affects both the fueling and HLS architectures.

The limiting case where propellant is only Earth-derived makes fueling on the lunar surface undesirable, whereas lunar-derived propellants are most easily used to fuel vehicles already on the surface. Whether or not lunar propellant is an option may impact the decision to leave a descent stage expended on the lunar surface, never to be refueled unless lunar propellant becomes available. If reusability is highly desired, but lunar propellant is not part of the trade space, a single element architecture that performs both descent and ascent and is refueled in orbit would more easily be reused than a split architecture where part of HLS needs to be replaced after every use.

Likewise, the level of reusability desired and level of ISRU available are interrelated when it comes to the refueling architecture. At one end of the spectrum, tankers could be manufactured and launched from Earth, fuel HLS once, then deorbit— each fueling mission requiring a new tanker. A slightly more reusable architecture could utilize a bus that intercepts tanks of propellant launched on reusable CLVs. The bus then ferries the propellant to HLS for refueling, and disposes the empty tanks. An additional step toward reusability could involve recovery of the tanks from Earth if they were equipped with a thermal protection system (TPS) and parachute or similar aerodynamic decelerator, however this extra mass would detract from the propellant payload. With

Lunar-derived propellant, the system can become fully reusable—either via fueling HLS on the surface or via a lander-tanker that can deliver propellant to orbit.

For now, a fully-reusable system is not in the near term (although technically feasible). More than likely, fueling architectures will have to start with Earth-derived propellant before transitioning to Lunar-derived propellants. The architectures assessed assumed that an architecture involving Earth propellants is necessary in the near term.

Steps to increase the number of reused elements in the architecture were examined, and options to evolve the architecture to utilize Lunar-derived propellants were a consideration.

2.3.2.1. Assumptions

The HLS architectures examined were used as inputs to the fueling architecture analysis with permission. The following assumptions were made:

1. Gateway is operational in NRHO and acts as an aggregation point for HLS

2. HLS is comprised of one of the following element combinations:

a. 2 element HLS: a transfer element (TE) and descent-ascent element

(DAE) or a descent element (DE) and ascent element (AE)

b. 3 element HLS: a TE, DE, and AE or an AE with two TEs

3. HLS is docked to Gateway before crew dock at Gateway via Orion

4. 4 Crew perform an intra-vehicular activity (IVA) from Orion to Gateway; 2 of

them IVA from Gateway to HLS

5. HLS undocks from Gateway bringing 2 crew to lunar surface while other 2

remain in Gateway

6. TE jettisoned once empty (baseline); TE loiters and is fueled for later missions as

an alternate architecture

7. 2 crew remain on lunar surface for 6.5 days and perform EVAs

8. 2 crew enters AE or DAE which returns them to Gateway

9. 4 crew enter Orion which undocks from Gateway and returns to Earth

10. AE (or DAE) remains docked to Gateway before being fueled by a commercially

launched tanker and is reused for an identical mission to the one described above

The HLS architectures considered were proposed by various internal subject matter experts (SMEs). All of the elements in each architecture were sized by an in-space propulsion SME using EXAMINE, an in-house tool based on detailed physical models and heritage hardware parametric models. The vehicle sizing models provided an estimate of the refueling propellant need for each architecture assessed. A diagram outlining the lander architecture was included in the Appendix.

Each architecture was analyzed by the author both qualitatively and quantitatively for potential reuse or partial reuse under the following different fueling scenarios:

1. Fully-expendable, commercially launched tankers (ZBO CFM)

2. A reusable bus delivers expendable tanks and provides ZBO CFM services

3. An augmented upper stage acts as a refueling element

a. Fast transfers to NRHO

b. Rendezvous with a tug in geosynchronous transfer orbit (GTO) which

spirals to TLI, then catches a lunar gravity assist to target a low-energy

transfer to NRHO

4. Varying degrees of refueling elements’ reuse were analyzed

Table 12. Propellant Need Assumptions: Human Lunar Lander

Total No. No. Elements Refuel Need per ID Elements Expended Propellant Type Mission (t)

1 3 2 LOx , LCH4 6.5

2 3 1 LOx , CH4 , LH2 21.1

3 2 1 LOx , LCH4 10.7

4 3 1 LOx , LCH4 14.8

5 3 1 NTO , MMH 14

Table 13. Launch vehicle capability assumptions

Launch Vehicle Assumed TLI Performance (t) Falcon Heavy (Fully Expendable) 18.7 Falcon Heavy (BR) 15.2 Falcon Heavy (B&CR) 11.5 New Glenn (CR) 14.9 Vulcan Centaur Heavy (CR) 11.3 OmegA Heavy (CR) 10.0

2.3.2.2. Campaign Trades

Two major HLS architectures will be traded: several two-element architectures and several three-element lander architectures. The performance of these architectures is not a consideration, but rather, the impact of these different architectures on fueling scenarios will be considered.

The human lunar campaign was assumed to consist of five HLS flights to the surface and back to Gateway. Growth to 50 and 500 missions was also considered. No lunar-derived propellants were assumed and the refueling of other systems besides HLS

68 was not considered (although most of the refueling elements could deliver Xenon to refuel Gateway’s PPE).

2.4. Useful Concepts

There are several ways to get to the moon: high-energy transfers (Hohmann), low- energy transfers (ballistic), and low-thrust transfers (spirals). The time of flight ranges from a few days, to a few months, or many months or more in each of the three cases respectively. For robotic spacecraft, all are part of the trade space, unless only chemical propellant is available, in which case only high and low-energy options exist. Low-thrust is exclusive to EP systems. High-energy, as the name suggests, uses more delta-v than low-energy. For tankers, maximizing delivered propellant per launch is a priority, so trading a longer transit time for a less “expensive” trajectory is desirable. Low-energy transfers are a particularly compelling option because they typically save more than ~400 m/s of delta-v (to an Earth-moon libration orbit), and have wide launch window options as well [40]. Low-energy transfers can also be done in tandem with one or even multiple lunar flybys to save even more on delta-v.

Low-energy transfers rely on throwing the spacecraft near the edge of the Earth’s sphere of influence where the sun can gradually raise the spacecraft’s perigee until the craft’s perigee is at a lunar distance, at which point the spacecraft can insert into the desired orbit “for-free”, or without any insertion burn. The first spacecraft to reach the moon on a low-energy transfer did so by accident. The Japanese mission Hiten originally was only supposed to send a probe to the moon, but after a communications array failure on the probe, mission designers frantically pulled together a new plan using a low-energy transfer to allow the entire Hiten spacecraft to reach the moon instead [40]. The first 69 spacecraft designed to use a low-energy transfer was GRAIL, followed by ARTEMIS which targeted an L1 and L2 Earth-moon Lagrange point.

Low-energy transfers are only possible in the three-body problem, in this case that is the Earth-Sun system (the third body being the spacecraft). A common approximation is to think about the circularly restricted three body problem (CRTBP), which simplifies the derivation. Consider a three body system where the third mass, 푚3, is negligibly small and has no impact on the behavior of the other two bodies. Recall the equation of motion for a mass in a rotating reference frame:

( ) ( ) ̈ ̇ 푟⃑ − 푟⃑1 푟⃑ − 푟⃑2 푟⃑ + 2휔⃑⃑⃑ ⨯ 푟⃑ = −휇1 3 − 휇2 3 − 휔⃑⃑⃑ ⨯ (휔⃑⃑⃑ ⨯ 푟⃑) ( 1) 휌1 휌2

Where:

2 2 2 2 휌1 = (푥 + 휇2) + 푦 + 푧 ( 2)

2 2 2 2 휌2 = (푥 + 휇1) + 푦 + 푧 ( 3)

In this problem, the Earth and Sun rotate about their barycenter with constant rotation rate 휔 = (0,0, 휔), but in our rotating reference frame, 푚1 and 푚2 are stationary.

Therefore, the position vectors 푟⃑1 and 푟⃑2 are constant. Normalizing the system such that

1 = 휇1 + 휇2 = 퐺푀 = 퐺(푚1 + 푚2) ( 4)

1 = 푟1 + 푟2 = 푅 ( 5)

And:

푟1 휇1 = ( 6) 푟2 휇2

It can be shown that 푟1 = 휇2 and 푟2 = 휇1, or rather 푟⃑1 = 휇2(−1, 0,0) and 푟⃑2 =

휇1(1, 0,0). Thus, by evaluating Equation 1, the equations of motion for the system become:

휕Ų 푥̈ − 2휔푦̇ = − 휕푥 ( 7)

휕Ų 푦̈ + 2휔푥̇ = − 휕푦 ( 8)

휕Ų 푧̈ = − 휕푧 ( 9)

where Ų represents the sum of the centrifugal and gravitational potentials:

휇 휇 휔2 Ų = − 1 − 2 − (푥2 + 푦2) ( 10.A) 휌1 휌2 2

By multiplying Equations 7—9 by 푥̇, 푦̇, and 푧̇ respectively and making use of the product rule, it can be shown that:

푑 1 [ (푥̇ 2 + 푦̇ 2 + 푧̇2) + Ų] = 0 ( 11.B) 푑푡 2

By integrating with respect to time we have:

퐶 = −2 Ų − 푣2 ( 12.C)

Where 푣2 = 푥̇ 2 + 푦̇ 2 + 푧̇2. This is in fact the Jacobi integral (and 퐶 is the Jacobi constant) [46].

Numerically solving the CRTBP yields 5 regions where 푣̇ = 푣 = 0 which are known as Lagrange points. These solutions are well documented in many other texts.

There are, however, infinitely many periodic and quasiperiodic solutions to the CRTBP.

If we linearize Equations 7-9 in the following way:

푥̈ ′ − 2푦̇ ′ − (1 + 2푐)푥′ = 0 ÿ ′ + 2ẋ ′ + (c − 1)y′ = 0} ( 13) 푧̈′ + 2푥̇ ′ + (푐 − 1)푦′ = 0

A characteristic equation linearly approximation of the 푥, 푦 motion is shown below:

′ 푥 = −푘퐴푦 cos(휆푡 + 휙) ′ 푦 = 퐴푦 sin(휆푡 + 휙) } ( 14) ′ 푧 = 퐴푧 sin(휈푡 + 휓)

Periodic motion results when the in plane and out of plane frequencies (휆, 휈) are equal. This condition is known as halo orbits [40]. Halo orbits are divided into two families: northern and southern. The former being the group of orbits where satellites spend the majority of their time above the Earth-Moon rotation plane. The halo orbits about E-M L1 and L2 can be reached by low-energy transfers. The subsequent mission design has been carefully studied by [40], [47], [48]. Low-energy transfers to specifically

NRHO were studied in detail by [49].

Ample mission data was available to assess architecture and vehicle concepts. For this reason, mission design was not an aspect of this investigation. The intended focus on vehicle and architecture design still warranted a comprehensive understanding of the orbital mechanics, however. The data gathered by [49] was instrumental in mission planning and the preliminary delta-v budgets generated for the various tankers that utilized low-energy transfers to rendezvous with NRHO located elements. This data will

72 also heavily impact scheduling, tracking, and the logistics network itself given the influence of Earth-Moon synodic period on mission parameters7.

3. Results and Discussion

3.1. Mars Hybrid Propulsion System Campaign Trades

3.1.1. Hypergolic Chemical Propellant Alternative

The Storable HPS full-factorial power analysis was examined. Crew trajectories were simulated to obtain the propellant needs for the matrix of Array and EP power levels. Using these needs, PropTracker was used to sweep through tanker mass fractions for each of the campaign cases. Not every case was evaluated, however. Only cases where the trajectories could close for every opportunity in the campaign were evaluated.

For some power levels sized, the storable HPS simply could not complete some of the missions. This was especially common for cases with lower array BOL power, but not exclusive.

Of the 32 power levels evaluated, payload PMF was varied from 0.5-1.0 (with an interval of 0.01), meaning in total, 1,600 distinct human campaign cases were simulated.

Out of all of these cases, most are insensitive to array BOL power. This is not unexpected. Although power available does impact the storable HPS performance, the main driver on propellant need, and thus tanker need, is the EP power. Plotted in Figure

11 is 18% of the 1,600 campaign cases evaluated.

7 Although mission parameters vary depending on departure epoch, mission parameters for lunar transfer trajectories repeat every Earth-moon synodic period (29.53059 Earth days). For this reason, once mission parameters are evaluated for a single synodic cycle, they do not need to be reevaluated for other cycles. 73

Figure 11. Storable HPS: Crew-Only Campaign Full Factorial.

Another common trend is the steep rise in tanker flights as EP power approached the lower bound. Not surprising, considering lower EP power generally requires more CP compensation, thus a higher total propellant need. For most cases, EP becomes inconsequential roughly around 400 kW and above. After increasing EP power beyond this point, most cases see no change in the number of tankers required. For this campaign specifically, this would suggest at least 350 kW EP power is desirable to be robust to changes in tanker mass fraction and array power. This does not imply overall vehicle robustness to other factors such as payload mass or launch opportunity, for instance.

Those would require separate analyses to understand.

In three instances, optimal sizes were apparent in the trade space analyzed. It is speculated therefore, that other cases have optimal points in this region as well, but it is not apparent from Figure 11. This finding verifies the trade space selected is appropriate,

74 since deltas to the design change are benign near the center of the trade space, yet begin growing significantly at some of the boundaries.

Although the tanker payload PMF is a major driver in the total number of fueling flights required, it would be inappropriate to assume other parameters are negligible in comparison. For the sake of completeness, a wide tanker payload PMF range was selected, although models indicate that for a tanker sized for a 15 t CLV, the payload

PMF would lean closer to the lower bound of the region considered. Secondly, another key caveat is that this analysis, unlike others in this text, neglected to consider any cargo flights in the campaign. Cargo trajectories, and subsequently their propellant needs, significantly differ from the crew trajectories considered. The departure and arrival destinations are the same, but as previously discussed, cargo deliveries are not constrained by time of flight, thus optimal cargo trajectories have different solutions. The data is not available to say for certain, but a reasonable hypothesis is that by including cargo flights, thus more HPS flights in general, the sensitivities present in the trade space would only be further amplified. For instance, any two given EP power levels might only differ by one additional tanker flight, but this gap would grow more significant with the inclusion of cargo flights.

3.1.2. Latitude Sensitivity

Propellant needs vary significantly from one Mars opportunity to another, meaning a robust transit vehicle design is important for being able to each Mars throughout the campaign. Placing any additional constraints on this goal only further confines the possible options. The capability to reach the same latitude each mission (so crew can return to a single stationary base on Mars) is one of those additional constraints.

A single vehicle robust enough to deliver crew and cargo to a range of latitude options

(the same one for each campaign case), is even more difficult.

Using trajectory data for the baseline HPS, delivery constraints were varied from -

50° to +50° latitude at increments of 10° [18], [26], [38], [48]. HPS is generally capable of delivering crew to a surprisingly wide range of latitudes (specific size varies by opportunity). For the baseline campaign, if a single latitude is inaccessible for one surface mission, that destination must be thrown out for the entire campaign (since the desire is to keep returning to the same surface location). The same must be done for cargo flights.

Cargo flights also have a surprisingly wide range of options, but unfortunately, the cargo has to be delivered to the same location as the crew, and these two regions only intersected at a few latitudes: -30°, -20°, and -10°. Although not nearly global access, this band stretches roughly 1,200 km north to south. The propellant needs for a campaign targeting each latitude in this range became the fueling scenario input.

Using the propellant needs for the range of accessible latitudes, the fueling scenarios were run for the baseline campaign. A 15 t to TLI CLV was assumed and tanker payload PMF was varied. Although crew and cargo propellant needs were available, the second type of cargo trajectories, “early” cargo, was unavailable for the range of latitudes under consideration. In the baseline, two cargo landers are delivered a

Mars opportunity before crew, but the lander that crew use to descend to the Mars surface must arrive with or before crew during the same opportunity as their surface stay. This constraint means the propellant needs for early and “free” cargo cases are not the same, so early landers had to be excluded from the analysis. Although an incomplete campaign, the analysis is still valuable because the relative differences between the latitudes can still

76 be examined with the information available. Secondly, two free cargo landers are needed for each early cargo lander, so the free cargo cases dominate the cargo propellant need nevertheless.

Since only three latitudes could be examined, the sensitivity is more difficult to analyze. Specifically, under the campaign considered (results may not apply to other campaigns and or opportunities), the fueling flights are relatively uninfluenced by latitude. Shifting the landing site either has no impact on the total number of tanker flights or only increases the number by one. Latitudes outside this 30 degree range cannot be reached by HPS at the specific opportunity and element age required (propellant need exceeds maximum tank capacity for at least one opportunity in the campaign). Thus, for achievable latitudes in the baseline campaign, there exists significant freedom in site selection with negligible fueling architecture impact.

Figure 12. HPS Campaign: Latitude Sensitivity on Fueling for 15 t CLV Tankers

3.1.3. Tanker Design Sensitivity

To fuel the baseline Mars campaign using expendable tankers will require procuring CLVs to launch the tankers. At today’s most competitive launch costs, this would be $6.75 billion FY19 USD8 assuming the baseline tanker mass fraction. The mean annual cost over the 17-year campaign would be roughly $397 million FY19 USD.

Of course, this cost is highly dependent on the CLV provider chosen and the mass

160

Falcon Heavy (RB&C) 140 Vulcan Centaur (RC) Ariane 6 Falcon Heavy (RB) 120 Vulcan Centaur Heavy (RC) New Glenn (RC) 100 Falcon Heavy (Expendable)

40 No. Commercially Launched Tankers Required Tankers Launched Commercially No. 20

0 0.5 0.6 0.7 0.8 0.9 1 Tanker Payload PMF

Figure 13. HIAD Lander Campaign: Tanker Design Sensitivity

8 This price reflects the best case scenario: the cheapest, most capable heavy lift CLV, Falcon Heavy (expendable). SpaceX CEO Elon Musk has stated the price for this launch configuration is $150 million [50]. 78 fraction of the tankers. Because of this level of uncertainty and the fact that launch costs themselves may change by the start of a Mars campaign, understanding the whole trade space is necessary. Figure 13 shows the sensitivity of tanker design and CLV choice on the total number of launches required. Note the nomenclature for the launch vehicles in the legend: recovered core stage (RC), recovered boosters (RB), and recovered boosters and core (RB&C). Each data point represents an entire campaign case run with

PropTracker. Fueling for 350 distinct scenarios was calculated in total. Representing the data in this manner was desirable because CLV provider availability, tanker design, and cost per launch is difficult to project decades from now. The performance impact of recovering CLV stages presently outweighs the cost savings from stage recovery and reuse. This is partly because a fully expendable Falcon Heavy is only $150 M versus $90

M USD for a Falcon Heavy with core and booster stage recovery. Granted, the second reason is that the estimated performance for the stage recovery configurations (taken from [51]) is likely conservative. It is customary to wrap these numbers with a manager’s reserve until the performance limits are well understood, then this margin can be retired.

With a campaign in the 2030s, there is no reason to include these reserves for a CLV that has already flown multiple times successfully.

3.1.4. Lander Trade

Figure 14. Lander Mass Sensitivity

As shown by Figure 14, lander mass does impact the campaign’s overall launches for HPS elements and associated fueling flights, as expected. Interestingly, the changes to the campaign caused by the added 14,000 kg to each lander are hardly impactful. Only one more HPS element is required to close the trajectory (by the end of life, HPS 1 cannot deliver the heavier lander, although it had no delivery problems for any missions at the beginning of life). Over the course of the entire campaign, the single extra element and the small number of additional tankers (exact number depends on which CLV is delivering the tanker), is relatively inconsequential. One note: the “mixed fleet” case assumed all three of the CLV providers are used at random. The distribution between each provider was as even as possible.

The campaign is still feasible with Mid L/D landers, which signifies two things:

First, HPS is robust to changes in both departure dates and payload requirements. Second, both Mid L/D and HIAD landers are still feasible options for a Mars campaign. That’s

80 not to say these capabilities will not have technical challenges throughout their development, but they both seem to be worth further consideration at this point in time and under the outlined assumptions of this study.

3.1.5. Tanker Thermal Control System Trade

Under the assumptions mentioned in Section 2, the mass delivered by the fast tanker architecture was determined based off the parametric tanker model. By taking this delivered mass and dividing by the total useful payload of the CLV, the PMF required for a slow tanker to deliver the same propellant mass was obtained. The results in Figure 15 can be interpreted as follows: given the hybrid propellant (capable of storing and transferring SEP and chemical propellant) tanker model assumed, and given a total expected boiloff percentage, the required PMF for a slow tanker to be at least as capable can be determined.

This result presents an interesting outcome: both the slow and fast tanker architectures appear comparable in the end. Although launching tankers on ballistic trajectories to save on delta-v may intuitively seem like the superior architecture, feasible fast delivery alternatives exist that may deliver comparable propellant loads in the end, with less technological development of ZBO TCS needed to accomplish this goal. Two caveats should be noted, however. This result is based off a first-order model, and is not meant to support or reject any specific technology, but rather, simply shows that further detailed investigation should be done to arrive at a more precise result if desired.

Secondly, this result cannot be generalized to other situations. The results were based off

81 a model of a specific tanker configuration, and a trade between two specific architectures.

Under different assumptions or situations, the results could be significantly different.

Figure 15. Cryogen Thermal Control Trade

3.1.6. Propellant Transfer Rate Trade

From an operations standpoint, refueling could either occur during the entire window where the elements are in Earth’s vicinity (excluding crewed periods), which maximizes the time for refueling, or refueling can be limited to NRHO for simplicity and or reduced risk. The latter significantly impacts the refueling window available. Given that the high-fidelity, end-to-end HPS trajectory revealed 8 months are required for departure operations and 6 months are required for arrival operations, the baseline campaign looks like the following:

Table 14. HPS Fueling Window

Earth Vicinity NRHO Opportunity (days) Stay Shakedown 731 491 2035 671 251 Table 6. Propellant 2039Thermophysical 882 Data Table 7.642 2039HPS Fueling Window517 277 2039 1,192 952 2039 473 53 2043 472 52 2043 472 52 2043 472 52 2043 472 52 2048 442 22 2048 442 22 2048 470 50 2048 545 305

Using the number of tankers required, the window available, and the amount of propellant required for transfer, the required transfer rate was determined for the two options:

Table 15. HPS Order of Magnitude Fluid Transfer Rate Requirements Fuel Only Fuel Anywhere Option in NRHO Uncrewed High 84 9.4 (kg/hr) Mean 29 5.4 (kg/hr)

Both of these transfer rates are within reasonable bounds, but having the capability to fuel anywhere is a major design driver for the fluid transfer system. Future work should investigate the feasibility of “anywhere” fueling.

3.2. Lunar

3.2.1. Fueling Architectures

A recurring architecture trade for all campaigns (lunar and Mars) will be the number of reusable elements in a fueling architecture. In order to consistently access the lunar surface, expendable (single-use) landers can be built for each trip (like Apollo), landers can be refueled and reused with expendable tankers, or landers can be refueled and reused with a reusable network of tankers. The recurring propellant demand of the landers drives the fueling architecture’s design.

3.2.2. Element Designs

3.2.2.1. Hybrid Common Bulkhead Tanker

The first tanker sized was the common bulkhead (CBH) monocoque design. The

CBH tank stores cryogens while a skirt connects a the CBH tank to a small SEP propellant tank and RCS tanks (SEP tank was assumed forward while one pair of RCS fuel and oxidizer tanks were assumed, and another pair aft). The barrel section of the monocoque tank assumes an isogrid structure for sizing purposes. Isogrids are a mass efficient, cutting edge structural design to withstand bending moments during launch.

The pressure-stabilized CBH design has significant heritage from its similarity to the

Centaur stage. ULA has invested significant research into adapting the Centaur to a long- lived cryogenic stage [52]–[54]. Significant progress has been made in multilayer insulation (MLI) for mitigating boiloff, even for Hydrogen. ULA has also invested in intra-vehicle fluids (IVF) technology which allows hydrogen boiloff to be repurposed to run combustion engines, act as a vapor cooling shield for tanks, and for other practical

84 applications. Although IVF and other technologies like autogenous pressurization will potentially make helium pressurant obsolete, they currently have no flight heritage.

Because the lunar campaign demands elements which are as near-term as possible, few low-TRL technologies were selected for the conceptual element sizing wherever possible.

The same trade studies could easily be repeated in the future with more new technologies.

The monocoque tank is sized to store cryogens in a densified state. Propellant densification is not a new concept. In fact, it has been studied in detail by NASA [55],

[56], and has been implemented by SpaceX on their Falcon 9 since 2016 [57], [58]. In addition to the launch vehicle benefits, propellant densification provides several key improvements for tanker designs. Near their freezing point or triple point, propellant density is considerably higher, meaning more propellant can occupy a given tank. See

Table 16 for density information. Secondly, the additional mass raises the “thermal inertia” of the tanker, meaning less boiloff occurs for a given heat leak. The added challenge to this approach is that tankers must “load-and-go” on the Launchpad, meaning propellants must be transferred to the tanker just before launch, just as SpaceX fuels their

Falcon 9 with densified liquid oxygen (DLO) at the last possible moment.

Table 16. Propellant Thermophysical Data

Hydrogen Oxygen Methane Krypton Xenon Normal Boiling Point (K) 20.3 90.2 112 120 165 Triple Point (K) 13.8 54.4 90.7 116 161 NBP Density (kg/m3) 70.8 1,140 422 2,417 2,942 Slush Density (kg/m3) 82.0 1,306 450 - -

Loading super-chilled propellants has caused problems for SpaceX in the past.

When the company first implemented DLO, they aborted several launches due to subsequent propellant loading problems on the pad. After 9 successful launches involving

DLO, the growing pains appeared to be over. However, in 2017, during a routine safety check just before a Falcon 9 launch, an explosion occurred in the second stage oxygen tank. The tank contains composite overwrap pressure vessel (COPV) helium pressurant bottles directly submerged within the oxygen tank. Failure investigators concluded a likely cause of the explosion was helium bottle buckling caused by DLO accumulation within the composite overwrap layers [59]. The company has stated the bottles were completely redesigned. DLO has not appeared to cause any additional problems since the

2017 explosion, but there is no doubt that propellant densification adds potential risk. In the case of SpaceX, those risks appear to be mitigated already, but propellant densification is still a potential risk worth tracking and a trade worth considering for a fueling architecture.

With added performance comes reduced launch reliability, SpaceX has learned.

This trade must remain a consideration especially for a fueling network that must be both highly reliable and cost effective. The risk is not constrained solely to the launch pad, however. To date, cryogen propellant transfer in-space has yet to be demonstrated, so the added potential risk and operational complexity of transferring densified propellants on orbit is a point of conjecture. The high uncertainty surrounding this capability warrants further analysis in order to understand the performance impact of the decision.

The SEP tank was sized using data based off of commercial off-the-shelf (COTS) tanks. SEP propellants have high densities which allow the tank size to be rather small compared to cryogens. The skirts were sized using a physical model approximating the necessary mass to support loads based on typical g-forces during launch for the geometry, mass, and center of gravity of the conceptual tanker. The inter-tank, forward, and aft

86 skirts were assumed to be lightweight composite material. The skirt must withstand bending loads (primarily during launch) between the two tanks. The center of mass of everything forward of the cryogen tank was estimated to obtain the moment arm of the bending load. The maximum bending load itself is simply the total mass forward of the cryogen tank multiplied by the maximum expected lateral acceleration during launch: 1.5 g (The skirt must also withstand an axial load of 5g). A safety factor of 1.4 was applied.

The tanker does not need a main propulsion system (MPS) because it performs no high delta-v maneuvers (injection, insertion, ascent, or descent burns). Since only low- energy burns are required (low-energy insertion, station keeping, trajectory correction maneuvers (TCM), phasing, rendezvous, docking, and escape burns), a reaction control system (RCS) is sufficient. The RCS utilizes NTO/MMH propellant (two tanks for fuel, two for oxidizer), two propellant feed systems, and 24 reaction control thrusters (12 forward, 12 aft). The tanks are sized based on the propellant needed for the delta-v budget, and the propellant feed system is sized based on the required system performance

(simple plumbing, valves, pump-fed system for mass estimate purposes only).

The fluid transfer system for the payload itself is sized using the same method as the RCS propellant feed system. The requirements differ, however. Insulated lines were assumed and electric driven pumps were included, but required flow rate was kept very low (<0.1 kg/s) to reduce the power requirements of the pumps. A faster transfer rate is theoretically, possible, but requires larger pumps and or higher pressure tanks.

The thermal environment is the most important input to sizing a spacecraft’s thermal management. The spacecraft can be subjected to wide swings in incident radiation throughout the mission, thus the effective ambient temperature varies

87 significantly. Conditions vary from launch (significant internal and external convection and radiation) to transit near Earth (high solar radiation and albedo), near Earth during eclipse (significant infrared), farther from Earth during direct sun (high solar radiation and low albedo), and farther from Earth during eclipse (less significant infrared).

Ultimately, the worst case drives the TCS mass requirements. For this architecture, the highest radiation flux the tanker will experience is direct sunlight just after fairing separation, where Earth’s albedo is highest (assuming launch occurs during local daytime). The heat leak calculated can be verified with this rough approximation which uses a 1-dimensional conduction model [60]:

퐾푇 ( 15) 퐹 = 푋

Next by assuming the tank is spherical with radius 푟, the boiloff rate can be approximated: 퐹 푅% = ( 16) 휌퐻푟

Granted, this worst case represents an off-nominal environment. For ~95% of the transit time of flight, the tanker would be far enough away that Earth’s albedo is negligible. Rough approximations of each of the aforementioned conditions were made to determine the worst case heat leak into the cryogen tanks. Even at worst case, it is well known that MLI can significantly reduce the heat leak. Equation 15 represents the theoretical optimal insulation thickness [60]:

퐾훥푇푡 푋opt = √ ( 17) 푀푓휌푖퐻

Once the MLI thickness was estimated with the model, the result was verified using

Equation 15. The MLI thickness became a heat leak estimation input, which was modeled and verified using Equation 13. Using the heat leak as an input, the active thermal management was sized.

The thermal control system (TCS) was modeled based partially on physical models and partially on Shuttle heritage TCS. Because the tanker is not crewed, no internal thermal control loop was necessary (toxicity to the crew was not a concern because the tanker is unmanned, and the campaign assumes fueling will only occur robotically). The loop assumes cold plates provide the thermal sink for the cryocoolers to acquire heat from the systems, a vacuum compressor heat pump (VCHP) operated with heat exchangers to another, hotter working fluid loop provide enough temperature rise for the radiators to effectively reject heat.

The avionics are all heritage, COTS components with nominal redundant systems.

The cabling was sized for the appropriate length of the vehicle and the appropriate hardware for AR&D was included. The communications must be capable of reaching

Earth from up to 2 million km away, roughly the maximum distance the spacecraft could reach on a low-energy transfer [40].

Power requirements were set by the needs of the spacecraft for avionics and thermal management. However, the driving power requirement was a function of the

TCS. Once total power needs were estimated, the solar arrays, boom, and related subsystems were sized based on the current latest and greatest, commercially available arrays. Batteries were driven by maximum eclipse duration. Assuming the 9:2 L2 NRHO is the loitering point for tankers, the maximum eclipse duration would be about 80

89 minutes with a frequency of only a few times per year. This short duration and low frequency is well within the capabilities of existing hardware9 [49].

3.2.2.2. Tanker Bus

As total campaign propellant need increases, the number of tanker flights increases to deliver the higher need. For campaigns requiring a large number of tankers, the fueling cost can become very significant. Instead of launching fully-expendable tankers (single-use), there is a point where it pays off to incorporate reusable refueling elements. The tanker bus concept was designed to help identify where that break-even point lies.

The expendable tankers require a lot of complex and expensive systems. The avionics capable of AR&D, the TCS capable of ZBO, pumps capable of transferring cryogens, and the high-power photovoltaics supporting these systems all must be state of the art in order to accomplish the mission. Consequently, the unit price of each tanker is rather high. To reduce the amount of “waste” from each expended tanker, a lot of these systems could be offloaded to a bus that stays in orbit and is reused for each new tanker.

The bus would provide power, ZBO cryocooling, AR&D, and the fluid transfer system for each tanker. Tankers in this architecture would be passively cooled and transfer to

NRHO on fast trajectories (4-5 days TOF).

In order to minimize the tankers as much as possible, a specific TCS was adopted.

Using a single loop reverse Brayton cycle, radiators, compressors, heat exchangers, and

9 This analysis did not predict the maximum eclipse duration during transit to NRHO. This could be a potential future task to verify it is less than the maximum eclipse in NRHO. 90 evaporators can all be integrated on the bus. The only TCS components necessary for the tanker would be fluid loop disconnects, plumbing integrated around the tank (broad area cooling), and MLI. The passive tanker has the same basic design as the actively cooled tanker, but with less thermal control and heat rejection, much lower power, reduced avionics, and a larger propulsion system and more RCS propellant (to perform the fast insertion burn).

Because the reusable bus is delivered on a BLT and does not perform high-energy burns (only a TCM, orbital maintenance, and AR&D are required), no MPS is required.

The RCS can also be fairly small because tankers can regularly resupply the bus with

RCS propellant (hypergolic). Assuming the bus is delivered by a heavy-class CLV, substantial mass margin (several metric tons) exists. This margin could be utilized to augment the mission and further reduce cost. A robotic manipulator launched with the bus could provide berthing to visiting tankers and vehicles. The option to berth would remove the burden of docking and reduce tanker cost. Although the common berthing mechanism (CBM) does not include fluid disconnects, engineers at NASA Ames

Research Center developed and patented a detailed design of a berthing mechanism which does support fluid connections [61].

Fueling is just part of the equation with reusable vehicle architectures. Batteries and arrays eventually degrade, subsystems fail and require repair or replacement. The robotic manipulator could also provide these services. Ultimately, the decision to berth or dock will come down to the operations and cost trades.

3.2.2.3. Long-Duration Upper Stage

Upper stages are fairly similar to tankers. They have large propellant tanks, feed systems, pressurant bottles and all the necessary systems of an expendable unmanned element. Many studies have previously investigated using modified upper stages like the

Centaur as a propellant depot. Significant analysis went into the necessary redesigns and found that passive thermal control, long-duration avionics and RCS, and settled pressure control were required key technologies. Even with liquid hydrogen, utilizing many passive cooling techniques significantly reduces boil-off by two orders of magnitude and provides ZBO liquid oxygen [53]. Clearly, upper stages have enough potential for use in refueling to warrant further consideration for at least first-order analysis.

The Centaur stage in particular was used as a starting point for several reasons: it has tremendous flight heritage as a heavy payload class upper stage (unlike the New

Glenn which has yet to fly); Centaur will remain flight ready in the near future on the

Atlas V and soon on the Vulcan Heavy (unlike the Delta cryogenic second stage which is being phased out); and finally because the Centuar uses cryogenic fuel, which must be delivered for the assumed campaign in this study (unlike the Falcon second stage which uses RP-1).

Two possible upper stage architectures were considered for this study. One involving upper stages transferring directly to NRHO (fast) then rendezvousing with a reusable bus. The second involves an upper stage launching to GTO10 then rendezvousing

10 GTO was selected as the intermediate node to avoid the intense radiation environment at lower altitudes. A SEP tug would be subjected to significant solar array degradation if it were to transfer slowly through the inner Van Allen Belts. Because the tug is reusable, spiraling through intense radiation for every single mission would significantly detract from the tug’s operational lifetime, rendering the architecture infeasible. 92 with a SEP tug that ferries the stage to NRHO. The tug performs the same functions as the bus (thermal control), but has more power and hall-thrusters to transfer to and from

NRHO with high efficiency low-thrust. Of course, upper stages consume much less propellant getting to the much lower C3 GTO node, but the drawback is that the tug is an additional system that must be procured. If the delta-v splits are chosen appropriately, a duplicate power and propulsion element (PPE) already being developed for Gateway could be procured and used as a tug, significantly reducing development cost. Of course, a SEP tug requires propellant as well, which means upper stages must be retrofitted with an additional tank and feed system for Xenon. Fortunately, Xenon is a high density, storable propellant which simplifies storage and transfer.

In order to analyze the upper stage architecture’s merit, the propellant load an upper stage could deliver to the staging point had to be estimated. CLV providers list approximate payload capabilities to various orbits and some other basic information about their launch vehicles. Unfortunately, a launch vehicle with effectively no payload

(aside from the minor modifications to the upper stage previously mentioned), has a significantly different ascent profile from the same vehicle pushing a payload to orbit.

Thus, the publicly available payload information on the various CLVs does not necessarily represent an accurate estimate of the propellant load remaining in an upper stage after launching with no payload. In order to determine a first order estimate of the

11,12 no-payload propellant remaining, CLV parametric models were utilized .

11 Here the parking orbit is where the upper stage first achieves orbit before performing the burn to reach its intended destination. The Atlas V Centaur assumed parking orbit is 185 km perigee, according to their publicly available launch service program (LSP) performance metrics. Other CLVs park in similar altitude LEOs. 12 CLV parametric models were not created by the author, but were used with permission for this analysis. 93

3.2.3. Performance

A matrix summarizing the refueling architectures evaluated is shown in Table 17.

This list of architectures is not exhaustive. The active tanker, passive tanker, bus, tug, and augmented upper stage correspond to the elements described in the previous sections.

Charts depicting the various architectures are shown in the Appendix.

Table 17. Propellant Delivered to NRHO to Delivered Propellant 17. Table

Figure 16. Expendable Tanker Mass Fraction Delta: Total Lunar Campaign Refueling Fights

An inert mass fraction growth was included in Figure 16. to underscore the importance of an efficiently designed tanker in the architecture. As the number of missions increases, and subsequently the propellant demand, a small delta in tanker inert mass can make an enormous impact on affordability across many missions. For 5 missions to the lunar surface, these are some different HLS architecture options considered. Each campaign was assumed to be refueled by the less efficient 7.1 t inert mass expendable tanker. Notice the baseline (far left) has the lowest potential for refueling (only 1/3 HLS elements refueled), thus has less tankers required.

Next, considering the architectures involving the (expendable) active tankers,

Figure 17 shows the number of commercially launched tankers required to supply the

HLS baseline campaign (AE refueling only). One important caveat here is the baseline

HLS only reuses a single element, the AE, so the TE and DE must be launched new for each mission. Another caveat is the underlying assumption that the tanker inert mass is static across all CLVs used. Because it was assumed too costly to design a unique tanker optimized for each of the many CLV providers available, the author chose to size a tanker

Figure 17. Refueling Flights: 7.1 t Tanker Inert Mass

to a CLV capable of 15 t payload to TLI. This assumption is important to remember because the smaller, less capable CLVs could deliver more propellant if they were not pushing an oversized tanker. Understanding the sensitivity of tanker gross mass on CLV fleets consisting of various payload capacities could be studied in the future.

The impact on HLS architecture was analyzed (Figure 18). The different architectures considered have different propellant choices, delta-v splits, and element sizes. The baseline requires the least number of tankers because only one out of three elements can be refueled (the rest must be relaunched each mission). All other architectures considered can refuel 1/2 or 2/3 elements, so their propellant demand is higher.

Expendable or partially expendable architectures do not scale well for large campaigns. The cumulative total number of stages and elements that would have to be expended to support the human exploration of the moon over 50 missions with AE and

TE refueling was determined for various tanker launch service providers (Figure 19.13

This specific HLS architecture was highlighted because NASA may evolve to an architecture much like this after the first few missions. Again, the DE must be expended because it was assumed that no lunar ISRU propellants are available.

Figure 18. 50 Missions: No. of Stages & Elements Expended: 7.1 t Inert Tanker, 3E TE Assisted Return – LOX/CH4

This tally does not include the launch of Gateway elements, logistics flights to support Gateway, cargo flights to the lunar surface, rover delivery, surface habitat and logistics delivery, and so on. Even with reuse of AE and TE, at minimum, 50 surface missions would be unaffordable with NASA’s current baseline architecture. With high enough flight rates (and thus a large enough propellant demand), there exists a point where it becomes more cost effective to trade a performance decrease for a fueling architecture where all or most of the refueling elements are reused.

13 This accounts for every single launch vehicle stage: core (if expended) boosters (if expended), upper stage, Orion capsule, DE, tanker, etc. 98

The principle of a refueling network is to save cost by reusing transportation systems like HLS. To benefit from this reuse, the refueling elements supplying the demand must either be considerably less expensive than the system being refueled or the refueling elements must be reused as well. As an analogy, a package delivery company would not throw away their trucks and buy new ones after delivering one shipment. Thus, meeting large propellant demands by trying to maximize the propellant load delivered per launch is not necessarily the best solution. Instead, one must investigate how to make the fueling network as cost effective as possible. The cost effective solution may or may not involve reusable refueling elements depending on the demand.

Expendable ZBO tankers are adequate for meeting a small propellant demand.

When scaled up to ~50 flights or more, they significantly drive the total campaign cost.

This was a key finding of the Mars exploration campaign. Throwing away tankers with complex avionics capable of AR&D, state-of-the-art large scale ZBO active CFM, large solar arrays, and the active fluid transfer system all significantly raise the tanker unit cost.

By simply off-loading these systems to an in-space bus that is reused for each tanker flight, we can still deliver close to the same propellant load as the “smart” tankers. Of course, the passive tankers will boil off propellant while they are disconnected from the bus, so they must transfer fast or suffer considerable boiloff. As a consequence, the passive tankers cannot take advantage of the free insertion provided by the ballistic transfer, so they must also include a larger propulsion system to reach NRHO. A trade was completed to compare the two tanker architectures, active and passive. The analysis accounted for no boiloff in the 4-day fast transit. This is a reasonable assumption if densified propellants and other advanced passive CFM are included such as improved

16.0 14.0 12.0 10.0 8.0 6.0 4.0 2.0 0.0 Falcon Falcon Falcon Vulcan New Glenn OmegA Heavy Heavy (BR) Heavy Centaur (CR) Heavy (CR) (Expendable) (B&CR) Heavy (CR) Passive Tanker + Reusable Bus Active Tanker

Figure 19.Passive Fast v. Active Slow Fueling Architecture: Delivered Propellant Per Launch

MLI. The results are in Figure 19. Interestingly enough, the passive tankers are able to deliver almost the same propellant load as the active tankers. This finding suggests having ZBO during lunar transit, specifically for LOX CH4 tankers, may not be a worthwhile investment, although ZBO certainly has powerful impacts in other architectures. ZBO CFM is still of course assumed to be a necessity for this architecture once the passive tanker reaches the bus in NRHO. The active and passive CFM tanker designs have inconsequential differences in performance for the tanker architectures assumed. This result assumes the passive tanker has negligible boiloff due to the advanced MLI, densified propellants, etc.

Rather than simplifying the tanker design (and still expending them), it may be considerably more cost effective to design a fueling architecture around full reuse. The following results will show how effective a reusable refueling network would actually be.

100

There are many ways an element can be reused. A tanker can launch, deliver, be recovered, then continue the same process repeatedly until the end of the tanker’s operational lifetime. Additionally, a refueling element can be reused by being repurposed.

The latter does not necessarily require the element return to its place of origin, but could remain at the destination and be used as a depot or habitation volume. Both types of reuse will be considered. The hope is that by maximizing reuse, the 400+ expended stages and elements required for the hypothetical 50 lunar missions (discussed previously) could instead be reduced to zero. This would, in turn, allow the campaign to be more affordable.

Figure 20. Propellant Delivered to NRHO: Upper Stage Stays in NRHO

Although several launch service providers have begun reusing stages or plan to reuse stages in their rockets, the upper stage is always expended. This stage is much harder to reuse because orbital reentry velocities require a large heat shield to prevent the stage from burning up. This added mass reduces the payload mass of the rocket. For now, companies like SpaceX have ditched the idea of recovering their upper stages because it is likely more cost effective than attempting recovery. To reuse a stage, it does not

101 necessarily have to be recovered, however. The stage can simply be repurposed. For instance, an upper stage like the proposed long-duration Centaur could launch without any payload to NRHO, transfer propellant to a lander, then stay in NRHO to be used as a depot at a later date. Once lunar ISRU propellants become available, the propellant could be transferred to the stages where they could be stored in orbit for later use.

Figure 21. Propellant Delivered to NRHO: SEP Tug; Upper Stage Stays in NRHO

A first order estimate of the propellant load various upper stages could supply to

NRHO is shown in Figure 21. One caveat is that the Falcon Heavy 2nd Stage uses RP-1 fuel. Since NASA has expressed no need for delivery of this fuel, the 2nd Stage can only deliver liquid oxygen without further modification (the RP-1 would not be transferred to any elements). This was accounted for in all of the analysis assuming a mixture ratio of

2.3. Upper stage delivery to NRHO assumes a mass penalty for Xenon propellant, tank, and disconnects, MLI, RCS, avionics, etc. to enable upper stage use for the low-energy transit to NRHO (with lunar gravity assist (LGA)). Each modification was sized based on the specific dimensions of the particular upper stage being used as a refueling element.

The tug was static, assumed to be 50 kW BOL arrays, 50% efficiency EP. A bus stacked

102 to the PPE was assumed to provide fluid transfer and ZBO CFM. Starship and New

Glenn were not assessed due to unacceptably long transit times for a tug of this size.

Because CLVs try to optimize their performance to the target orbit of their customers, they typically have better performance to orbits like GTO rather than TLI.

NASA’s recent procurement of the Gateway Power and Propulsion Element (PPE) also suggests the agency is interested in gaining experience flying high-power hall thrusters.

Another fueling architecture assessed bridged these two ideas. By launching upper stages to GTO14, CLVs can deliver more propellant. Additionally, a tug can take advantage of

SEP’s high efficiency to further increase propellant delivered. The SEP tug can potentially be a duplicate of the PPE, which could minimize development cost. However, upper stages would have to be further modified with a small SEP tank and fluid disconnect to fuel the tug. Also, the upper stage would require a bus to provide CFM, but the power can be provided by the tug. Even with these masses accounted for, the tug architecture can deliver up to 100% more propellant compared to the upper stage architecture without a tug.

Additionally, this tug architecture has other benefits. Instead of repurposing upper stages in NRHO, we will now assume it is more desirable to return the upper stages empty after they transfer propellants in NRHO. While none of the upper stages considered can deliver significant propellant to NRHO and return to LEO all under their

14 GTO was also selected as the staging point for the following reasons: low-power SEP tugs have extremely long transit times if departing from LEO (they are already several hundreds of days if departing from GTO), and the intense radiation environment roughly between LEO and GEO would significantly degrade solar arrays if they were in this environment for long durations. 103

Figure 22. Propellant Delivered to NRHO: SEP Tug; Upper Stage Returns Propulsively own propulsion, the tug makes this possible for all the heavy CLVs, aside from New

Glenn and Starship.

Another way to increase the propellant delivered to NRHO without the use of a tug would be “topping off” upper stages in LEO with distributed launches so the upper stage performs the TLI burn at maximum propellant load. Several CLV providers are currently pursuing this idea, namely the SpaceX Super Heavy Starship and the ULA

Vulcan Centaur. Top offs have the better performance, but require multiple launches to deliver a single upper stage to NRHO.

All of the expendable cases were compared on the basis of propellant delivered normalized by number of launches to understand the added value of distributed launch cases. In all of these cases compared, the refueling element itself does not return to Earth

(although the tug returns to GTO). According to the data shown in Figure 23, the highest propellant load delivered per launch, not surprisingly, is by the largest proposed rocket, the Super Heavy Starship. For the other CLVs, the Tug + upper stage architecture has the highest delivery per launch and the upper stage by itself has the lowest. Surprisingly, introducing reusable elements actually increases delivered payload. Also, since the

104

Figure 23.Refueling Architecture Comparison: Propellant Delivered to NRHO Per Launch: Refueling Element Stays in NRHO tankers were all sized to be about 15 t gross mass or greater, topping off Vulcan or

OmegA upper stages (or even launching a single upper stage) actually has better performance than using these CLVs to launch slightly oversized tankers. For these architectures specifically (where the refueling element is left at NRHO), introducing reusable elements like tugs or topping off stages actually increases propellant delivered on a per launch basis.

4. Conclusions and Future Work

A fueling network approach enabled by CLVs is a feasible alternative to propellant depots. The architecture offers adaptive, on-demand delivery for variable mission frequencies and destinations. Another key figure of merit is the significantly lower launch cost than SLS propellant delivery—assets delivered to NRHO would cost roughly $9,800/kg FY19 USD on a fully-expendable Falcon Heavy compared to

$77,000/kg FY19 USD on an SLS Block 1 Cargo. Launching propellant on CLVs rather

105 than SLS cargo offers lower launch costs, a scalable solution, and the flexibility to launch propellant from many CLV providers.

Strategies to further increase performance of commercially launched fueling architectures were found. Passive CFM tankers can deliver 96% of the propellant mass per launch as active CFM tankers. Long-duration upper stages deliver more propellant to

NRHO than ZBO tankers (launched on the respective CLV) if, on average, tanker inert mass exceeds 51% of the launch vehicle usable payload. Topping-off CLV upper stages with propellant in a LEO parking orbit can deliver more propellant per launch than upper stages launched directly to NRHO. Starship Super Heavy and the New Glenn, if topped- off in LEO then launched fast to NRHO, could deliver 230% and 130% more propellant per launch (including top-off launches) than active CFM tankers launched on a fully- expendable Falcon Heavy to NRHO (on low-energy transfers). A 50 kW SEP tug improves long-duration upper stage performance by a mean of 180% for Falcon Heavy,

Vulcan Centaur Heavy, and OmegA Heavy compared to long-duration upper stages with no tug. Based on the relative improvements found in this first-order investigation, these architectures should be kept in the trade space for future campaigns.

The second useful outcome was the viability of PropTracker for automating propellant tracking in order to feed into transportation system and campaign level analysis. PropTracker evaluates fueling architectures for over 1,000 unique campaigns per hour, compared to the ~5 per hour previously determined manually. Thus,

PropTracker widens the potential trade space by rapidly evaluating different scenarios, enabling more informed trade studies and high-level decisions.

106

In order to better leverage the current capabilities of PropTracker for new applications like exploration campaigns involving ISRU or for commercial use, several updates to PropTracker are suggested. Mainly, improving scheduling and timing simulations would complement the existing tracking features. Currently, the user must define their own launch vehicle capabilities or use the default models. Connecting

PropTracker to a more standardized and external launch vehicle model would establish commonality, offer greater flexibility, and increase automation. Automatically reading output from vehicle sizing models as input for the tanker models would allow faster and more architectural trade studies to be completed than currently possible with manually setting up inputs. Allowing a high degree of flexibility for the user to plug PropTracker into other specialized models such as boiloff models would broaden the potential applications.

Another potential use of PropTracker would be campaigns that leverage ISRU propellant production. The production could occur anywhere including the Moon, NEAs, or Mars. Essentially no modifications to PropTracker would be necessary for this application, aside from importing a new vehicle model for the desired mission. The model that currently represents a tanker could instead represent a rover that gathers resources for processing or a lander that delivers propellant to orbit.

1. Study failed delivery scenarios (failure and risk analysis)

2. Implement loading and scheduling optimization

a. Based on RBO H2 and ZBO O2

3. LOX/H2 network for updated NTR

107

Human exploration is enabled by in-space fueling. Of course, certain capabilities like ISRU propellant production cannot occur without some form of propellant transfer.

This capability, as already alluded to, will also enable new markets for commercial space operations. Space is still inaccessible to many companies and nations because it requires them to build their own satellite and find a launch service. The single largest contribution to the operational lifetime of the satellite is the propellant onboard. If instead of a single use, satellites could be refueled and reused hundreds of times, they would not only become more cost-effective, but space debris generation would be rapidly curbed.

Organizations could also lease time on reusable satellites rather than launch their own— further increasing access to space and reducing cost. Planned constellations like Starlink, consisting of tens of thousands of satellites, will also benefit from a carefully designed refueling architecture. The methods developed and used to study human exploration architectures in this paper could be used to conceptually design commercial refueling architectures.

PropTracker has high relevance to an application with much higher current commercial demand than human exploration: satellite refueling. Thousands of satellites currently orbit Earth and not a single “gas-station” is currently able to provide any of them with even a single reuse. There is a high demand for satellite refueling because it could extend the mission for satellites that cost millions of dollars to develop and launch as well as curb space debris build up from “single-use” satellites. PropTracker could easily be used for architecture trade studies by essentially running the program logic in the reverse direction. Instead of small tankers supplying a much larger transit vehicle with propellant, a comparatively large depot would siphon off propellant to smaller

108 satellites requiring a refuel. The element propellant needs would instead represent initial depot propellant loads and each fueling event would involve propellant transfer from the depot to the visiting satellite. With very slight modifications to the current version of

PropTracker, trade studies such as point designs or sensitivity analyses could be run for theoretically any number of depots or satellites. One could quickly size a fueling network for constellations currently being built-up such as Starlink. Likely key outputs of interest that could be studied with PropTracker would be the right combination of launch vehicle choice, parking orbit choice, depot size, depot quantity, propellant transfer efficiency, propellant transfer time, constellation size, and satellite propulsion system size.

NASA is planning to construct Gateway, an outpost in NRHO which could be utilized for staging, outfitting, assembling, or fueling missions to the lunar surface, Mars, or near Earth asteroids (NEAs). If more than one of these missions or campaigns overlaps in operation schedule, tracking assets like propellant becomes even more necessary.

Assuming Gateway were to provide CFM, power, thermal management, orbital maintenance, or a docking port to elements being fueled or waiting for fuel, tracking will inevitably be necessary. Knowing how long and when elements need specific Gateway assets is crucial for coordinating other missions with other elements that may need the same assets soon before or thereafter.

Campaign trades for an overlapping lunar and Mars exploration could be investigated. As an example, a Lunar and a Mars campaign could be assumed to have overlapping mission operations where Gateway is leveraged as a fueling location for both campaigns. The PropTracker results could generated based on the Mars HPS campaign and a notional human lander system (HLS) for lunar surface access.

109

PropTracker is not restricted to tracking only propellant. The code is generic in the sense that it only passes doubles between elements, thus those doubles can represent masses of cargo and logistics in addition to propellant. In fact, some vehicles that provide resupply, like Progress, are capable of offloading both cargo and propellant. Using

PropTracker for such applications would require little to no modification depending on the specific application. This use for PropTracker is recognized, but will be a part of future work.

Finally, understanding what is the most cost effective fueling network solution as a function of demand would be highly valuable. Right now, it is unknown if or when a fueling network with reusable elements could become less expensive than an expendable delivery approach. Based on the successes of companies like SpaceX, the trend appears to be in favor of more reuse, but further analysis is required to confirm this hypothesis.

Ultimately, the most affordable fueling network possible is necessary in order to lower the entry barrier of space exploration for smaller agencies and companies with ambitions in space. Eventually, the demand may become high enough that fueling networks could become profitable, just like fuel transportation services on Earth. The path that lies ahead depends on our ability to make forward-thinking and informed decisions today.

5. Appendix

The HLS being refueled had 5 notional architectures. The first one served as a baseline—shown below:

110

Element Element Architecture

. Human Landing System: Baseline 3 Landing System: . Human

111

The refueling architectures to deliver propellant to HLS are summarized below. For each case, the refueling element is either a unique tanker sized for the mission, or a CLV upper

stage modified for long-duration:

. Refueling .Element Refueling (CLV Upper orStage Tanker)

112

ling Element ling (Either UpperCLV orStage Tanker) Bus + Reusable

Refue

- A

113

SEP TugSEP + UpperStage:CLV NRHOEnds Upper Stage in

114

SEP Tug + CLV Upper Stage: Upper Stage Returns to Returns LEO Stage TugStage: Upper Upper SEP + CLV

- A

115

6. References

[1] M. J. Neufeld, “von Braun and the lunar-orbit rendezvous decision: finding a way to

go to the moon,” Acta Astronautica, vol. 63, no. 1–4, pp. 540–550, 2008.

[2] S. A. Jefferies and R. G. Merrill, “Viability of a Reusable In-Space Transportation

System,” in AIAA SPACE 2015 Conference and Exposition, Pasadena, California,

2015.

[3] “Satellite Servicing Projects Division,” SSCO. [Online]. Available:

https://sspd.gsfc.nasa.gov/. [Accessed: 26-Oct-2019].

[4] D. Chato, “Experimentation for the Maturation of Deep Space Cryogenic Refueling

Technologies.” NASA Glenn Research Center, 2008.

[5] R. E. Coffey, “The application of depot transfer/rendezvous in space based cyclic

orbit missions,” 1990.

[6] D. Chato, “Technologies for refueling spacecraft on-orbit,” in Space 2000

Conference and Exposition, Long Beach,CA,U.S.A., 2000.

[7] M. Garcia, “About the Russian Progress Spacecraft,” NASA, 25-Sep-2017. [Online].

Available:

http://www.nasa.gov/mission_pages/station/structure/elements/progress_about.html.

[Accessed: 14-Sep-2019].

[8] “Progress Cargo Craft Docked to ISS suffers Fluid Leak – Progress MS |

Spaceflight101.”.

[9] “FISO telecon07-09s archive list.” [Online]. Available:

http://fiso.spiritastro.net/archivelist07-09.htm. [Accessed: 14-Sep-2019].

[10] “Restore-L Factsheet.” NASA Goddard Spaceflight Center, 2018.

116

[11] “Robotic Refueling Mission 3 Can’t Perform Cryogenic Fuel Transfer,” in

Parabolic Arc, 22 April, 2019, [Accessed: 14-Sep-2019].

[12] “Orbit Fab demonstrates satellite refueling technology on ISS”, Space News, 18

June, 2019 [Accessed: 14-Sep-2019].

[13] “Orbit Fab raises $3M to make orbital refueling easier, cheaper and more

accessible,” TechCrunch.

[14] S. Potter, “NASA Announces Industry Partnerships to Advance Moon, Mars

Technology,” NASA, 30-Jul-2019. [Online]. Available: http://www.nasa.gov/press-

release/nasa-announces-us-industry-partnerships-to-advance-moon-mars-

technology. [Accessed: 24-Sep-2019].

[15] B. G. Drake, S. J. Hoffman, and D. W. Beaty, “Human exploration of Mars, design

reference architecture 5.0,” in 2010 IEEE Aerospace Conference, 2010, pp. 1–24.

[16] B. G. Drake and D. Watts Kevin, “Human exploration of Mars design reference

architecture 5.0, addendum# 2,” 2014.

[17] D. A. Craig, P. Troutman, and N. Herrmann, “Pioneering space through the

evolvable Mars campaign,” in AIAA Space 2015 Conference and Exposition, 2015,

p. 4409.

[18] R. G. Merrill, P. Chai, C. A. Jones, D. R. Komar, and M. Qu, “An Integrated Hybrid

Transportation Architecture for Human Mars Expeditions,” in AIAA SPACE 2015

Conference and Exposition, 2015, p. 4442.

[19] H. Price, J. Baker, and F. Naderi, “A minimal architecture for human journeys to

Mars,” New Space, vol. 3, no. 2, pp. 73–81, 2015.

117

[20] K. E. Goodliff, B. Mattfeld, C. Stromgren, H. Shyface, and W. Cirillo, “Comparison

of Human Exploration Architecture and Campaign Approaches,” in AIAA SPACE

2015 Conference and Exposition, 2015, p. 4413.

[21] D. A. Smith, “Space Launch System (SLS) Mission Planner’s Guide,” 2018.

[22] Statement of James Bridenstine Administrator National Aeronautics and Space

Administration before the Subcommittee on Commerce, Justice, Science, and

Related Agencies Committee on Appropriations United States Senate. 2019.

[23] G. Cates, C. Stromgren, D. Arney, W. Cirillo, and K. Goodliff, “International

human mission to Mars: Analyzing a conceptual launch and assembly campaign,” in

2014 IEEE Aerospace Conference, 2014, pp. 1–18.

[24] V. S. Reddy, “The spacex effect,” New Space, vol. 6, no. 2, pp. 125–134, 2018.

[25] “NASA Strategic Plan 2018,” p. 64.

[26] P. Chai, R. G. Merrill, and M. Qu, “Mars Hybrid Propulsion System Trajectory

Analysis, Part I: Crew Missions,” in AIAA SPACE 2015 Conference and Exposition,

2015, p. 4443.

[27] S.-L. Chen and Y.-Y. Chen, “Design and implementation of a global logistic

tracking system based on SaaS cloud computing infrastructure,” J. Syst. Manag. Sci,

vol. 1, pp. 85–96, 2011.

[28] P. Nicholls et al., “Logistics system for automating transportation of goods,” May-

1997.

[29] A. Shamsuzzoha and P. T. Helo, “Real-time tracking and tracing system: Potentials

for the logistics network,” in Proceedings of the 2011 international conference on

industrial engineering and operations management, 2011, pp. 22–24.

118

[30] V. Vavrík, M. Gregor, and P. Grznár, “Computer simulation as a tool for the

optimization of logistics using automated guided vehicles,” Procedia engineering,

vol. 192, pp. 923–928, 2017.

[31] S. Wang and Q. Meng, “Sailing speed optimization for container ships in a liner

shipping network,” Transportation Research Part E: Logistics and Transportation

Review, vol. 48, no. 3, pp. 701–714, 2012.

[32] W. Cirillo, C. Stromgren, and G. Cates, “Risk Analysis of On-Orbit Spacecraft

Refueling Concepts,” in AIAA SPACE 2010 Conference & Exposition, Anaheim,

California, 2010.

[33] “NASA Systems Engineering Handbook,” p. 297.

[34] “System of Systems,” College of Engineering - Purdue University. [Online].

Available: https://engineering.purdue.edu/Engr/Research/Initiatives/Archive/SoS.

[Accessed: 14-Nov-2019].

[35] W. Heinemun, “FUNDAMENTAL TECHNIQUES OF WEIGHT ESTIMATING

AND FORECASTING FOR ADVANCED MANNED SPACECRAFT AND

SPACE STATIONS,” p. 52.

[36] A. Ketsdever, “Spacecraft Design and Sizing,” University of Colorado Colorado

Springs Department of Mechanical and Aerosapce Engineering.

[37] P. R. Chai, R. G. Merrill, K. G. Pfrang, and M. Qu, “Hybrid Transportation System

Integrated Trajectory Design and Optimization for Mars Landing Site

Accessibility,” AIAA Propulsion and Energy 2019 Forum, Aug. 2019.

119

[38] P. Chai, R. G. Merrill, M. Qu, P. D. Kessler, and R. T. Joyce, “Sensitivity Analysis

of Hybrid Propulsion Transportation System for Human Mars Expeditions,” in

AIAA SPACE and Astronautics Forum and Exposition, Orlando, FL, 2017.

[39] P. Chai, R. G. Merrill, and M. Qu, “Mars Hybrid Propulsion System Trajectory

Analysis, Part II: Cargo Missions,” in AIAA SPACE 2015 Conference and

Exposition, Pasadena, California, 2015.

[40] J. S. Parker and R. L. Anderson, Low-Energy Lunar Trajectory Design:

Parker/Low-Energy. Hoboken, NJ, USA: John Wiley & Sons, Inc., 2014.

[41] J. Jarrige, P.-Q. Elias, F. Cannat, and D. Packan, “Performance comparison of an

ECR plasma thruster using argon and xenon as propellant gas,” in Proceedings of

the 33rd International Electric Propulsion Conference, 2013, pp. 2013–420.

[42] M. R. Nakles, W. A. Hargus Jr, J. J. Delgado, and R. L. Corey, “A Performance

Comparison of Xenon and Krypton Propellant on an SPT-100 Hall Thruster

(Preprint),” AIR FORCE RESEARCH LAB EDWARDS AFB CA, 2011.

[43] M. Nakles, W. Hargus, J. Delgado, and R. Corey, “A Performance and Plume

Comparison of Xenon and Krypton Propellant on the SPT-100,” in 48th

AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, 2012, p. 4116.

[44] D. Jacobson and D. Manzella, “50 kW class krypton Hall thruster performance,” in

39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 2003, p.

4550.

[45] R. R. Sostaric, C. J. Cerimele, E. A. Robertson, and J. A. Garcia, “A rigid mid lift-

to-drag ratio approach to human Mars entry, descent, and landing,” in AIAA

Guidance, Navigation, and Control Conference, 2017, p. 1898.

120

[46] “Co-Rotating Frame.” [Online]. Available:

http://farside.ph.utexas.edu/teaching/336k/Newtonhtml/node123.html. [Accessed:

28-Sep-2019].

[47] J. S. Parker, “Targeting Low-Energy Ballistic Lunar Transfers,” J of Astronaut Sci,

vol. 58, no. 3, pp. 311–334, Jul. 2011.

[48] J. S. Parker, Aas 06-132 Families of Low-Energy Lunar Halo Transfers. .

[49] “White Paper: Gateway Destination Orbit Model: A Continuous 15 Year NRHO

Reference Trajectory,” National Aeronautics and Space Administration (NASA),

Aug. 2019.

[50] M. Sheetz, “Elon Musk says the new SpaceX Falcon Heavy rocket crushes its

competition on cost,” CNBC, 12-Feb-2018. [Online]. Available:

https://www.cnbc.com/2018/02/12/elon-musk-spacex-falcon-heavy-costs-150-

million-at-most.html. [Accessed: 20-Oct-2019].

[51] M. Carney, “Launch Vehicle Performance Website.” NASA Launch Services

Program.

[52] B. Kutter et al., “Atlas Centaur Extensibility to Long-Duration In-Space

Applications,” in Space 2005, Long Beach, California, 2005.

[53] G. Szatkowski, M. Holguin, and B. Kutter, “Centaur Extensibility for Long

Duration,” in Space 2006, San Jose, California, 2006.

[54] R. Honour, R. Kwas, G. O’Neil, and B. Kutter, “Thermal Optimization and

Assessment of a Long Duration Cryogenic Propellant Depot,” in 50th AIAA

Aerospace Sciences Meeting including the New Horizons Forum and Aerospace

Exposition, Nashville, Tennessee, 2012.

121

[55] R. O. Ewart and R. H. Dergance, “Cryogenic Propellant Densification Study.”

NASA Lewis Research Center on Behalf of Martin Marietta Corporation, Nov-

1978.

[56] M. M. Fazah, “STS Propellant Densification Feasibility Study Data Book.” NASA

Marshall Space Flight Center, Sep-1994.

[57] T. Fernholz, “The ‘super chill’ reason SpaceX keeps aborting launches,” Quartz.

[Online]. Available: https://qz.com/627430/the-super-chill-reason-spacex-keeps-

aborting-launches/. [Accessed: 29-Sep-2019].

[58] “Safety panel considers SpaceX ‘load-and-go’ fueling approach viable,”

SpaceNews.com, 18-May-2018. [Online]. Available: https://spacenews.com/safety-

panel-considers-spacex-load-and-go-fueling-approach-viable/. [Accessed: 29-Sep-

2019].

[59] “SpaceX completes failure probe, plans new launch.” [Online]. Available:

https://www.cbsnews.com/news/spacex-completes-failure-probe-plans-january-8-

launch/. [Accessed: 29-Sep-2019].

[60] Spacecraft Mass Estimation Relationships and Engine Data. NASA Johnson Space

Center Advanced Propgrams Office, 1988.

[61] M. Cohen, “SPACE STATION ARCHITECTURE, MODULE, BERTHING HUB,

SHELL ASSEMBY, BERTHING MECHANISM & UTILITY CONNECTION

CHANNEL,” N84-22612, 1984.

122