LMUILA Loyola Marymount University Entitlement Spending Is Encroaching on the • External budget pressures crowding out defense Defense Budget

Spending as a Percentage of GDP spending 1o;; .-----.-----.----~ r----;~~~~ • Years of underinvestment • Two protracted wars • Continuing economic slowdown • Spending on entitlement programs • Interest on debt

1965 1970 1975 1900 1985 1990 1995 200:J 2005 2010 Shifting Balance in the Defense Budget (est) Source: U.S. Office of Manageinent and B'4let. HiitDOOJITobles, Budget of die Uniti!d Stales Defense Resources, in Billions of 20 IO Dollars GM:!mmert. Rsw/)001 20 I I {W.tlli1gton, D.C: US. Goverrvnent Printing Office, 2010). Tilies 8.4, 85, and l 0.1, at hrrp://v.\1w.wNrd1ruse.ga../anb/budgetify2009/pdflhistpdf $14 7 ······" Operations and (Ma.rd12, 20IO). Maintenance • Internal pressures and shifting priorities crowding out

$178········ Personnel modernization and procurement • Escalating costs of personnel compensation,

$ / /5·· ···· ·· Procurement retirement benefits, and health programs • Shrinking economies of scale and build rates

$50 ----l $5 4 ···· ···· Research, • Secretary of Defense forced to make major Development,Test. 2010 2028 and Evaluation changes to new and existing programs Sou rce: Congressional Budget Office,"Long-Term Implications ofthe Rscal Year 20 IO Defense Budget;· January 20 10, at htr.p:lhvww.cbo.god ftpdoes/lOB>txldocl Ofi52101 -25-FYDP.pdf(Ma.-ch 17.2010). LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 2 • The Defense Industry could be considered an oligopoly • Few large, competing firms respond to each other's actions • Produce a homogeneous product; Cost/Price discriminating variable

• Defense systems large and complex • Many inputs to the Cost/Price variable • Space example: Cost Savings Opportunities • Launch Segment Price fixed by Launch Providers No practical re-use options Lowest price LV selected

• Ground Segment Use existing Ground assets Re-use existing SW

• Space Segment Lower cost by re-using qualified S/C components Re-use lowers NRE costs

Costs can be lowered by re-using existing assets and reducing development costs by building additional components to qualified requirements.

LMUILA Loyola Marymount University Luis Aguilera- Spring 2012 3 • Objective: To develop a set of requirements applicable to components across multiple programs and missions; and through re-use minimize NRE costs.

• Rationale: - At time of topic selection, author involved in similar employer­ funded I R&D effort. - Approach was to envelope different types of requirements (i.e. Performance, Environmental, Mission Assurance, etc) to meet multiple programs and missions. - Requirements development assigned to functional expertise RE's. - Programs committing to "multi-platform" components pay NRE development costs. - Future programs would benefit from reduction in NRE costs. - Follow-on to a previous, highly successful effort 15 years prior

LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 4 • Set of re-usable requirements for use across multiple-missions and platforms • Spacecraft-level designers • Locate components on the SIC with respect to LV interface to attenuate the environmental effect on a component to within capabilities • Design component mounting methods to attenuate environmental effects to within maximum capabilities • Component-level designers • Design implications on selection of internal compenents, layout, etc • Design standardization to maximum environmental effects • Designs limited by capability of internal components

LMUILA Loyola Marymount University Luis Aguilera- Spring 2012 5 • The project is about requirements development and re-use. • Considers design/requirements re-use as an input to cost reduction. • Environmental requirements are interface requirements between systems: - Spacecraft / Ground Processing - Spacecraft/ Launch Vehicle - Spacecraft / Space • Basic spacecraft design drivers and component layout considerations - Launch Vehicle • Mass properties • Static and dynamic envelopes - Spacecraft • Mass properties • Thermal balance • Functional location • Integration • Physical accommodation

LMUILA Loyola Marymount University Luis Aguilera- Spring 2012 6 • Re-use as a Lean principle • Lean design seeks opportunities to eliminate waste during the design phase • Many sources of process waste in product development • "Re-use" focuses on Re-invention waste • Same problem is solved repeatedly • Existing solutions/designs are not utilized • Advantages are lower costs, faster development and lower risks • Develop component and requirements/performance standard re-use to lower costs • Specific-use = development N RE costs each time ~ Cl :f • Re-use= development NRE costs once i • Subsequent builds incur production costs only ·~..c: 3----c: • Barriers to Re-use ~ (!) • Corporate culture, Program-centric mentality • Re-use successes Test Cases, Vectors, Results

• Propulsion system components The Re-use Pyramid LMUILA Loyola Marymount University Luis Aguilera- Spring 2012 7 • Approach for this project follows a portion of a similar effort done in the early 1990s for Propulsion System components • Multiple-step process over 5 years

Evaluate Select Develop, ID needs Define and Develop EQ Maintain and vendor bids vendors and test and and bound specs and update and refine negotiate qualify components requirements standards specs requirements agreements products

• Accomplishments: - 12 subcontracted components standardized, design and tested for multiple programs application > 2 dozen piece parts and assembly hardware standardized - 10 standardized propulsion modules assembled from standardized components and piece parts Streamlined test/ acceptance procedures and QA inspections for standardized components Established long term agreements with Suppliers Standardized piece parts in Standard Stock Standardized products have been good for 15 years LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 8 SIC Pre-launch Environments Thermal A/C & GN2 EMC (RFIESD) Contamination I Cleanliness

SIC Launch Environments Design Load Factors Acoustics Sinusoidal Vibration Random Vibration 1-----_.. Component-level ~---1 .....1-- .. Specification Shock Thermal De pressurization Contamination EMI

SIC Operational Environments CMEslSolar Flares Geomagnetic Storms High Energy Particles Electromagnetic Radiation Space Plasma Atomic Environment Micrometeorites

LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 9 Evaluate Select Develop, ID needs Define and Develop EQ Maintain and vendor bids vendors and test and and bound specs and update and refine negotiate qualify components requirements standards specs requirements agreements products

• Some of the most severe environments a SIC can experience occur during lift-off/launch • Developing a robust set of enveloping requirements begins with the selection of available Launch Vehicles (Appendix A) • Previous efforts compared program defined environments • A survey approach was taken to compare environments from Mission Planner's Guides source documents • Enveloped environments according to a "90°/o rule" • Data points above the 95o/o or below the 5% would be considered "outliers" and treated on an individual basis • Not all LVs compared may have environments as severe as the enveloping curves but they will be "blanketed" with the approach

LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 10 • Mission Radiation requirements more severe than those used to envelope the Standardization/Re-use effort • Pressure Transducer particularly sensitive to the radiation environment • Options: - Develop a program/mission-specific component requiring entire development/qualification effort • Significant cost • Schedule penalty - Use an existing component with a work-around plan • Selected a standardized Pressure Transducer • Met increased radiation requirement by using a tantalum sleeve

For "outlying" requirements (beyond the enveloping requirements of a re­ use/standardization component), minimally invasive engineering solutions should always be explored before deciding on new development programs.

LMUILA Loyola Marymount University Luis Aguilera- Spring 2012 11 • Thermal • Electromagnetic Compatibility (EMC) - Launch Vehicle Intentional/Unintentional RF Emmisions - Launch Range Electromagnetic Environment Limitation - Electrostatic Discharge - Lightning Mitigation • Contamination Control/Cleanliness • Conditioned Air /GN2 purge • Humidity

LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 12 Design Limit Load Envelopes • The combined spacecraft accelerations are a function of launch vehicle 12 - V - 40Z, 50Z characteristics as well as spacecraft 10 - - 4YZ, 5YZ dynamic characteristics. 8 - Atlas V- < 7000 lbs -- - Delta IV- M/M+(4,2) • Design Load Factors are for use in 6 preliminary design of primary structure. 1;~ ~ I - IV- 4 M+(5,2)/M+(5,4) C) - ~ ~ Delta IV - Heavy • Axial acceleration is determined by the -II) 2 C: - Falcon9 vehicle thrust history and drag. 0 :;:::. '"~ - ctl 0 • - zenit3SL s.. • Maximum lateral acceleration is ~ ~.!t !~ Q) -2 - c., - - primarily determined by wind gust, ;i -4 - Envelope (Liquid engine gimbal maneuvers, first stage Motor) I - max -ctl -6 engine shutdowns, and other short­ ~ - Minotaur IV- max duration events. -8 - Taurus • -10 Positive axial value indicates a - Envelope (Solid Rocket Motor compressive net-center of gravity -12 acceleration -10 -5 0 5 10 Lateral Accelerations (g) • Negative value indicates tension. • Two envelopes are presented because a single envelope is too conservative and impractical. • Higher accelerations in the Taurus may be typical of solid-rocket based launch vehicles. • Red envelope representative of the Taurus. • Green envelope representative of the envelope for the liquid-rocket based launch vehicles. LlVfOTLA oyo a arymount niversity Luis Aguilera - Spring 2012 13 Acoustic Environment During Liftoff and Flight • Acoustic energy is the primary source AtlasV40z --+-- Atlas V41z (4-m ) (4-m Payload Fairing) ·- Atlas V42z ~ AtlasV43z of vibration to a space launch vehicle. (4-m Payload Fairing) (4-m Payload Fairing) -a-Atlas V54z Atlas VSSz (5-m Payload Fairing) (5-m Payload Fairing) • Transmitted to S/C by fluctuating --e-- At las V4zz (0.85 m from PLF vents) --+-- Atlas V4zz (0.61 m from 4-m PLF vents) pressures within the LV fairing; -"jg_ - Atlas V4zz (0.30 m from 4-m PLF vents) -- Delta IV Medium/Medium+(4,2} 4-m Composite Fairing

~ --Delta IV Medium+ (5,2 and 5,4) 5-m Composite Fairing -- Delta JV Heavy 5-m Metallic Fairing impinge directly on exposed SIC Q) O? surfaces, inducing vibration in high c:'!135 e gain antennae, solar panels and I r::a other components having a large ~ 1125 ratio of area-to-mass. Q) ...I • Fluctuating external pressure field e::::, ti) causes an oscillatory response of the et11115 CL rocket structure, ultimately transmitted "C C: through the S/C attachment ring in the ::::, ~105 form of random vibration. • A single enveloping requirement would be impractical and overly-conservative 95 for smaller payload components. l'O 10 10000 ° Frequency (Hz) lOOO • Red envelope represents the enveloping acoustic requirement for the large LV class (i.e. Delta IV, Atlas V, and 3L). • envelope represent the enveloping acoustic requirement for the small LV class (i.e., Taurus, & IV, Falcon 1 and Pegasus). LMUILA Loyola Marymount University Luis Aguilera- Spring 2012 14 - - Low-ShodlEnYlllope Enveloping LV Shock Levels - AlluV Type APLA • SIC are subjected to pyroshock 10000 ...... AIIUV Type 8119<4PLA(NearCBOO) -----· AllasV TypeB1194PLA(l0-J50degrees lromCBOO) environments which are LV-induced --- · AJj nV Type 01666PLA but can also be self-induced.

1000 • Pyroshock events are a result of the firing of an explosive device, usually for the purpose of initiating or performing a mechanical action.

- 1.11nolaur l Separa WngShc,;J,; • Spacecraft separation events, the ·········• l.llno\ilur l Non-5e pa 1i111"f9Shod: - Mlnolaur r,t Pll. to l V release of propulsion system safing , ...... , 1.tlnota urfV LVloP/L (StandardCli!mpBand) - - - - - · l.llnotaur rv LVloPA..(llg htBand) devices and deployable launch lock - - - · l.llnol.:lurfV 10 LVloP/1.(Noo•PA.S epa,ati(l'ISRS) !,,,·· - Peg n us ,,, .. Ba seolP/L devices are typical such mechanical __ - Taurus ,,, .. J Base olPII. - Zenl13SL 1194VSPAF(S1CSepar.1t:on•Radl1 INcl1) action. T ······• Zlml13Sl 1194VSPAF(SIC5eparat:on•Alcla1Alcl1) · ····· Ze ni13SL I r 1194VS PAF(AllOOW ILV ShockEvenls• RadlalAlol s) • Pyroshock events can also occur in - - - · Zenl\3SL 1194VSPAF(AIIOlhu llVShockEv1nls-Ax!alAxls) 100 1000 100QQ- - Zenl\3SL 1666PAF(A1 ShockEvenlsundA.1Cei;) flight during engine firing for Frequency (Hz) -- Zem!JSL 702•fl02•GEMPAf(S1CSepa r.1t on-AI Aice1) correction.

• Units should be designed to survive a shock spectrum that considers all shock sources. • Shock requirements may apply to in-plane or normal-to-mounting plane. • Moving units farther away from shock sources can provide some attenuation of the shock spectrum __by, virtue of distance or seQaration by number of structural j,_o_in_ts_.______. LMU LA Loyola Marymount University Luis Aguilera - Spring 2012 15 • Electronic devices are usually enclosed within a controlled thermal environment in the spacecraft interior. • This thermal environment can be tightly controlled (5-10°C) through a combination of passive heat conduction through a high-thermal capacity base plate and active heating elements. • Components on spacecraft exterior can be exposed to temperatures from -120 to +150°C. • Thermal requirements are extended to a broader range to guarantee components will function after accidental loss of power or overheating.

Maximum/Minimum Predicted Temperatures (MPT) 60 l .~ J. j \. ., l. -j l Test & Duration (Thermal Vacuum - TV) Cycles 50 MJ - -· - ·- - - f f Acceptance Env MPT + min range 14 J l Protoqualification Ace ± 5°C 27 40 I I \ ' ' , Qualification Ace ± 10°C 27 1'- J I \, Test & Duration (Thermal Vacuum - non-electrical} ' , Acceptance MPT 1 I ' Protoqualification Ace ± 5°C 3 ~ 20 :I " Qualification Ace ± 10°C 6 -~ Test & Duration (Thermal Cycle - TC) [ 10 ., Acceptance Env MPT + min range 14 E j Protoqualification Ace ± 5°C 27 ~ 0 \ ~ Qualification Ace ± 10°c 27 ., ., l Test & Duration (Combined TV and TC) -10 :, I ' Acceptance Env MPT + min range 4 TV+ 10TC Protoqualification Ace ± 5°C 4 TV+ 23TC -20 I \ ~ Qualification Ace ± 10°C 4 TV+ 23TC -30 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21

LMUILA Loyola M;;~~unt University Luis Aguilera- Spring 2012 16 Worst Case Electric Field Emissions S/C Electric Field Radiation Impingement on LV 170 180 1 I 1 160 I 11111 150

140 I I 1 1 11 11 I I I I II I I I J 120 I 11 Ill I I 11 I 11' :!!. - Taurus ,, - AtlasV400 100 I I I I I 111 111 I I II I II 11 11 I Ill 1111 ~ - At lasV - At lasVSOO u I II I I 11 II I I I - Delta lV-4m ...... , Delta JV tJ ! 80 - Zen it (Sea l aunch) ,, - Delta IV-Sm C: ~ Minotaur I .. .D - Zen it (Sealaunch) I II -.-Minotaur IV 3 60 ~ - Falcon9 --Falconl z.. 40 I I

70 20 I 11 I I

0 II I 111 I I I 50 I I I 1.Et04 1. E+OS 1.E+06 1.E+07 1.E+08 1.Et09 1.E+10 1. Et11 1.E+04 1.E+OS 1.E+OG 1.E+07 1.E+OS 1.E+09 1.E+lO 1.E+ll l.E+12 1.E+13

Frequency (H z) Fre quency (Hz)

• Electromagnetic Compatibility (EMC) for a system or component requires that : • It does not cause interference with other systems or equipment • It is not susceptible to emissions from other systems, equipment or electrical environments. • It does no cause interference within itself that can cause the system or equipment to malfunction or behave in an undesirable manner. • Graph on left showns maximum LV emissions a component must be to withstand. • Grapb oo r:igbt sbnws maximum emissions allowed fr:om a SLC tbat a L~ can witbstaod ____ LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 17 Results: SIC Operational Environments

• Space operational environment is difficult to characterize for any given m1ss1on. • Unknowns in mapping it and in processes that generate it. • Environment also changes with time, including launch timeframe, mission duration, and orbit characteristics. • Natural space environments - Solar-radiation events • Electromagnetic radiation • High-energy particles • Low-Medium energy particles - Atomic environment - Micromeorites • Man-made space environments - - Active spacecraft

LMU ILA Loyola Marymount University Luis Aguilera- Spring 2012 18 + ELECTRON MAGNETIC HIGH ENERGY LOW-MEDIUM ENERGY RADIATION PARTICLES PARTICLES

ARRIVAL: IMMEDIATELY ARRIVAL: 15 MIN TO FEW HOURS ARRIVAL: 2 - 4 DAYS DURATION: 1-2 HOURS DURATION: DAYS DURATION: DAYS

X-RAYS, EUV, GEOMAGNETIC EVENTS RADIO BURSTS STORMS

SATCOMINTERFERENCE SATELLITE DISORIENTATION SPACECRAFT CHARGING RADAR INTERFERENCE SPACECRAFT DAMAGE SATELLITE DRAG SHORTWAVE RADIO FADES RADIO BLACKOUTS POWER BLACKOUTS

LMUJLA Loyola Marymount University Luis Aguilera- Spring 2012 19 Space Environment Effect on Space System Mitigation Strategies Space Debris Puncture, Degradation, Collision Threat mitigated by reducing amount of space debris - Deorbit and re-orbiting at EOL - Passivation and discharging battery at EOL Small objects - shielding techniques Large objects - calculate collision probability (monitoring and tracking) ' Colllision avoidance

Active S/C Unintentional: frequency disturbances Development of space situational awareness systems Intentional: frequency disturbances, - Monitor S/C, identify their missions and predict possible threats electromagnetic , partial/total damage Orbital Perturbations Short-periodic, Long-periodic, Secular changes in orbital Natural threats more un-predictable than man-made threats elements - Fundamental phenomenon not fully understood CMEs/Solar Flare Atmospheric drag, Material degradation, Electronic component failure Threats highly dependant on mission duration, orbital altitude, and time within naturally occurring cycles Geomagnetic Storms Atmospheric drag, Electronic component failure High Energy Particles Deep dielectric charging, Single-event effects, Memory Modeling phenomenon's in future during design phase protects SIC from failure, Degraded sensor performance harsh space environment Electromagnetic Radiation Thermal effects, Ionization, Material degradation Real time, space environment monitoring integrated with on-board SOH Space Plasma Surface charging, Induced electric current, Arc discharge sensors can allow SIC to auto-adapt to changing conditions

Atomic Environment Chemical reaction with surface, Ionization, Atmospheric Switching off S/C during solar storms reduces damage from increased drag radiation dose. Micrometeorites Puncture Source: Space Environment Threats and Their Impact on Spacecrafts m Near Earth (Singh, Anafar and Nicolas) The Space Environment is difficult to characterize (establishing an enveloping requirement) due to it's dependence on the mission duration, orbital parameters and the time within naturally occurring space cycles.

LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 20 • This environment enveloping approach can be used on many spacecraft components. • Payload components such as optical and science instruments are impractical because they are so mission-specific and few are built. • Approach is better suited on spacecraft components such as actuators, drive units and other card based or slice-based components • Cost savings are realized as soon as a component is reused. • The first unit incurs the NRE costs • Each additional unit only incurs the RE costs • These savings can be used to pass on to the Customer in subsequent proposal bids • Design re-use is an ambitious undertaking • Can be a daunting task if scope is not clearly defined • Too broad can lead to 'chasing one's tail' and failure to converge on a usable set of requirements • Too narrow can yield none or insignificant savings

LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 21

First Launch 2002 2002 2004 2006 2010 1990 1994 1999 2000 2010 Stages 2 2 2 2 2 3-4 3 2 4-5 4

12,500 9,390 {27,558) (20,702) Capacity to LEO - 22,977 420 10,450 440 1,458 580 1,735 to to N/A kg (lb) (50,646) {924) {23,050) (970) (3,214) {1,300} (3,889} 20,520 13,360 (45,238) {29,440)

7,095 7,510 {15,642) {16,550} Capacity to 550- 22,560 420 8,560 190 1,054 331 to to N/A N/A kg (lb) (49,740) (924) {18,870) (420} (2,324) (730) 14,096 11,300 {31,076) (24,920)

4,750 4,541 (10,450) {10,012} Capacity to GTO - 13,399 4,540 430 6,180 to to N/A N/A N/A N/A kg (lb) (29,540) {10,000} (950) (13,624) 8,900 7,020 (19,580) {15,470)

VAFB CCAFS CCAFS CCAFS Kwajalein CCAFS Various VAFB Launch Sites VAFB Pacific Ocean VAFB VAFB VAFB CCAFS Kwajalein (air-launched) MARS KLC

Note: ULA is no longer producing the Delta II and is therefore excluded from the survey. Three launches of the Delta II are scheduled for 2011 with five launch vehicles remaining (unassigned). (2011 U.S. Commercial Space Transportation Developments and Concepts: Vehicles, Technologies, and Spaceports)

LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 23 Appendix B: Environmental Test Levels and Durations

Test Qualification Protoquallflcatlon Acceptance Source Static Load Level 1.25 x Limit Load 1.25 x Limit Load 1.0 x Limit Load Duration GSFC-STD-7000 1 minute 30 seconds 30 seconds 5 cycles@ full level 5 cycles@ full level 5 cycles@ fu ll level Acoustics Level 6 dB above Acceptance 3 dB above Acceptance Envelope of MPE & MIL-STD-1540E Duration 3 minutes 2 minutes 1 minute Random Vibration Level 6 dB above Acceptance 3 dB above Acceptance Envelope of MPE and M IL-STD-1540E Duration 3 minutes each of 3 axes 2 minutes each of 3 axes 1 minutes each of 3 axes

Sine Vibration 1.25 x Limit Level 1.25 x Limit Level Limit Level GSFC-STD-7000 2 act/min 4 act/min 4 act/min Shock Level 6 dB above Acceptance 3 dB above Acceptance Maximum predicted MIL-STD-1540E Duration 3 X both directions 2 X both directions 1 X both directions 3 orthogonal axes 3 orthogonal axes 3 orthogonal axes Pressure Level As specified in Appendix A As specified in Appendix A 1.1 x MEOP (structures) MIL-STD-1540E Duration Thermal Vacuum Level ±lO"C beyond Acceptance ±5°C beyond Acceptance Envelope of MPT and MIL-STD-1540E (-24 to 61 "C) Duration 27 cycles 27 cycles 14 cycles Thermal Vacuum Level ±l0°C beyond Acceptance ±S"C beyond Acceptance Maximum predicted MIL-STD-1540E Duration 6 cycles 3 cycles 1 cycle Thermal Cycle Level ±lO"C beyond Acceptance ±5°C beyond Acceptance Envelope of MPT and MIL-STD-1540E (-24 to 61 "C) Duration 27 cycles 27 cycles 14 cycles Combined Thermal Level ±lO"C beyond Acceptance ±S°C beyond Acceptance Envelope of MPT and MIL-STD-1540E (-24 to 61 "C) Duration 4 TVs+ 23 TCs 4 TVs+ 23 TCs 4 TVs + 10 TCs EMC Level 12 dB 6dB 6dB MIL-STD-1540E Duration Same as Acceptance Same as Acceptance 20 min. @ each SV

LMUI LA Loyola Marymount University Luis Aguilera - Spring 2012 24 Sine Vibration {Axial & Lateral) 4.0 • Sinusoidal vibration is used to

3.5 simulate the effects of significant flight environment launch 3.0 transients. - Atlas V (Axial) 2.5 - Delta IV (Axial) • These transients typically produce - Taurus (Axial)

- Zenit (Axial) the dominant loading on primary - Lv Envelo pe (Axi al) and secondary structure and many - Atlas V (Lateral) 1.5 of the larger subsystems and - Delta IV (Lateral) - Taurus (Lateral) assemblies. 1.0 I - Zenit (Lateral) • Only current method of adequately -.. - - - LV Envelope (Lateral) 0.5 exciting the lower frequency dynamic modes, particularly those 0.0 below 40 Hz. Frequency (Hz)

• Sine vibration levels are dependent on the type of structure to which the component is attached, the local attachment stiffness, the distance from the spacecraft separation plane, and the component's mass, size and stiffness; impractical to specify a single enveloping requirement . • Actual practice is to specify multiple vibration spectrums based on above parameters for design __guidelines. . LMU!LA Loyola Marymount University Luis Aguilera - Spring 2012 25 • Random vibration environment LV Random Vibration Envelope simulates the response of the - Minotaur! ------+ (P/VEnviro nmenl) spacecraft bus structure to the - Minotaur IV (Structure-Born MPE) required acoustic spectrum. ---l- I - Minotaur IV I (Full Spectrum Vibroacoustic MPE) • Base-input random vibration - Pegasus (P/Linterface - X-a:,cJs) analysis and testing is appropriate -- Pegasus (P/ linterface - Y-a1ds) for units that are relatively compact ...... Pegasus N ....,: (P/ linterface -Z-axis) and do not in themselves have a lot :l9 0.0100 +----l"'------'-,,L----,,<----,---,1---1---'-"~·-"----'-----'---~~ - Taurus Q (P/llnte rface) ~ of surface area that would be - zenit3SL (P/llnterface) directly driven by the acoustic

- Falconl (Mll-ST0-1540E Min RV Spectrum sound pressure. - 1000 lbs P/L) - Falconl (1000 lbs P/ L) • Random vibration input occurs over - MIL-ST0-1540E Min RV Spectrum (P/L< 50 lbs) a broad frequency range, from

- Combined LV Random Vibration Envelope about 10 Hz to 2000 Hz.

10 100 1000 10000 • A minimum level of vibration is Frequency (Hz) necessary to ferret out existing workmanship defects and potential failures. • Difficult to characterize a single enveloping Random Vibration requirement because they are affected by location from base vibration input, mass and size of the unit and the adjacent components on the panel. • Actual practice specifies several Random Vibration environments based on the above parameters. LMU LA Loyola Marymount University Luis Aguilera - Spring 2012 26 • Fairing Internal Pressure Environment Payload compartment pressure and

120.0 ------~ --AtlasV431(4mFairing)-min depressurization rates are a functions o - Atlas V431 (4m Fairing)· max the PLF design and trajectory. 110.0 -1----1---~'---l--+--+--+---+---+----+---+----+--t---+---+----i---< - Atlas V431 (Sm Fairing)· min

- Atlas V431 (Sm Fairing)· max ~illk--~ l------1-1--+-!---+--l---+--+--+--+---+---+--+--I • As a launch vehicle ascends through the - Atlas V Design - Early

90.0 --"-'-·'-+- ____,_--l-_.,____----1--1---1----1-----1---i--1---1---l--l - Atlas V Design -Transonic atmosphere, venting occurs through the aft - Atlas V Design-Late

M section of the fairing and other leak paths 0. .l< 80.0 -l---+---\.l.\'l."i-----',-1----\" ...j--+-!---+--l---+--+--+--+---+---+--+--I --Delta IV Medium · min 1! - Delta IV Medium - max in the vehicle. ~ £ 70.0 +--+--J--1-.-\\\AA--l-¥.-+~ --+--l---+--+--+--+---+--+--+--I-I - Delta IV M+ 14,2) · min ~ - Delta IV M+ 14,2) · max • Knowledge of depressurization rates and ~ E 60.0 -l----l---~1--\'Ml,~~-\'-+-!---+--1---+--+--+--+---+--+--I Ji - Delta IV M+ (5,2) · min 0. venting design concepts are important E -l----l----l----1\-\~\.'lM,l--+l+-l~ --+--l---+--+--+--+---+----+--I --Delta IV M+ (5,2) · max u0 so.a .,, - Delta IV M+ (5,4) · min because improper venting of high of ~ >M 40.0 !l.\.--'\cl-.\----\i------'lt..--1-!---+--l---+--+--+--+--I --Delta IV M+ (5,4) · max 0. - Delta IV Heavy (Composite PLF) · max pressure could lead to compression and 30.0 thus inducing structural collapse , ~ Sea launch-max

-++-Sea Launch - min penetration of sealing and interference - Falcon! with component function - Fairing Pressure Envelope - max ~--at;;;l:::j:~ - Fairing Pressure Envelope - min • Conversely, low pressure could result in 0 10 20 30 40 so 60 70 80 90 100 110 120 130 140 150 160 170

FlightTime (sec) expansion and fracture of a container, explosive expansion, loss of mechanical strength, insulation break down and arc­ over as well as loss of mechanical strength and alteration of electrical properties.

LMUJLA Loyola Marymount University Luis Aguilera - Spring 2012 27 Appendix B: Space Environment Threats

Orbits Likelihood Severity (km above Earth's surface) Noc = # of registered occurences until now (effect on SIC and/or onboard Instruments and/or whole mission (SOIM)) L LO . Minor SOIM degradation (LEO): 200 to 1,200 Very Seldom (VS): N c < 3 0 ow ( ). ( .,I.. 0%-25% effective life time)

. Partial SOIM degradation Seldom(SE): 3 < N c < 5 Medium Earth Orbit (MEO): 3,000 < Apogee < 30,000 0 Medium (ME): ( .,I.. 25%-50% effective life time)

. . Significant SOIM failure nchronous Earth Orbit (GEO): - 36,000 Often (OF): 6 < N0 c < 8 Critical (CR): ( .,I.. 50%-75% effective life time)

. Complete SOIM failure Very Often (VO): 8 < N c 0 High (HI): ( .,I.. 75%-100% effective life time)

LEO MEO GEO Threat Likelihood Severity Threat Likelihood Severity Threat Likelihood Severity Space Debris SE LO, ME, CR, HI Space Debris vs LO, ME, CR, HI Space Debris vs LO, ME, CR, HI

Unintentional: LO Unintentional: LO Unintentional: LO Active SIC vs Active SIC vs Active SIC vs Intentional: CR, HI Intentional: CR, HI Intentional: CR, HI

Orbital Orbital Orbital VO LO,ME VO LO VO LO Perturbations Perturbations Perturbations

CMEslSolar Flare SE HI CMEslSolar Flare SE HI CMEslSolar Flare SE HI

Geomagnetic Geomagnetic Geomagnetic SE HI SE HI SE HI Storms Storms Storms

High Energy High Energy High Energy OF CR OF CR OF CR Particles Particles Particles

Electromagnetic Electromagnetic Electromagnetic OF CR OF CR OF CR Radiation Radiation Radiation

Space Plasma OF ME Space Plasma OF ME Space Plasma OF ME

Atomic Atomic Atomic OF ME OF LO OF LO Environment Environment Environment

Micrometeorites OF CR Micrometeorites OF CR Micrometeorites OF CR

Source: Space Environment Threats and Their Impact on Spacecrafts in Near Earth Orbits (Singh, Ariafar and Nicolas) LMUILA Loyola Marymount University Luis Aguilera - Spring 2012 28 Appendix M Difficulties in Enveloping Operational Environments: Radiation

Plasma interference Space hazard Spacecraft charging Single-event effects Total radiation dose Surface degradation with communications

Trapped Solar Trapped Solar Ion Wave Specific cause Surface Internal Cosmic rays o •erosion Scintillation radiation particle radiation particle sputtering refraction

LEO <60°

LEO >60°

MEO

GPS GTO I GEO HEO - Interplanetary - -- Not Important Relevant applicable ~ LEO <60°-low Earth orbit, less than 60 degrees in cl ination; LEO >60°-low Earth orbit, more than 60 degrees inclin ation; MEO-; GPS-Global Positioning System satellite orbit; Source: GTO-geosynchronous transfer orbit; GEO-; 1. J.E. Mazur, "An Overview of the Space Radiation Environment" Cross/ink, Summer HEO-hig hly ell iptica l orbit; 2003, www.aero.org o•- atomic oxygen LMUILA Loyola Marymount University Luis Aguilera- Spring 2012 29