A Study of CPS Stages for Missions beyond LEO Final Distribution

Size: px

Start display at page:

Download "A Study of CPS Stages for Missions beyond LEO Final Distribution"

Cory Washington
6 years ago
Views:

1 A Study of CPS Stages for Missions beyond LEO Final Distribution 31 May 2012 Atlanta, GA Mark Schaffer Senior Aerospace Engineer, Advanced Concepts Group

2 Introduction Contents 1. Introduction to SpaceWorks 2. Study Overview 3. Parametric Sensitivity Analysis 4. Application to CPS Design 5. Architecture Solution 6. Summary and Conclusions 7. Appendix Background 6 month study from June through December 2011 Sponsored by United Launch Alliance Performed by SpaceWorks with support from United Launch Alliance 2

3 Introduction to SpaceWorks 3

with headquarters in Atlanta, GA and facilities in Washington, DC and Huntsville, AL Founded in 2000 as a spin-off from

5000 list of fastest growing, privately-held small businesses ranked #39 in Engineering Over $2M annual revenues from

4 SpaceWorks Enterprises, Inc. (SEI) Overview Atlanta, GA Washington, DC Huntsville, AL Aerospace engineering services and space systems analysis firm with headquarters in Atlanta, GA and facilities in Washington, DC and Huntsville, AL Founded in 2000 as a spin-off from the Georgia Institute of Technology Named to 2011 Inc list of fastest growing, privately-held small businesses ranked #39 in Engineering Over $2M annual revenues from sales of engineering services and software SBA small-business classification DCAA-audited and approved accounting system Three primary operating divisions: Engineering, Commercial, and Software Two subsidiary companies: Generation Orbit Launch Services, Inc. Terminal Velocity Aerospace, LLC. 4

Corporate Organization SEI Affiliates/SMEs Norm Brown Carl Ehrlich Rich Reinert William Rothschild

Olds Chief Financial Officer Rodney Bonsu Business Manager Ashley Russ Office Manager John Bradford,

5 Corporate Organization SEI Affiliates/SMEs Norm Brown Carl Ehrlich Rich Reinert William Rothschild SpaceWorks Enterprises, Inc. (SEI) John R. Olds, Ph.D., P.E. Chief Executive Officer Melinda S. Olds Chief Financial Officer Rodney Bonsu Business Manager Ashley Russ Office Manager John Bradford, Ph.D. Division President Jon Wallace Division President Advanced Concepts Group (ACG) Brad St. Germain, Ph.D. Director of Advanced Concepts Dominic DePasquale, CEO Engineering Economics Group (EEG) Nicole Martin, Ph.D. Director of Engineering Economics Space Media Group (SMG) Mark Elwood Director of Space Media 5

SpaceWorks Personnel High percentage of technical staff Degrees

engineering, media arts and animation, and marketing Advanced degree

(25%) Affiliates with extensive industry experience including

6 SpaceWorks Personnel High percentage of technical staff Degrees including aerospace engineering, mechanical engineering, industrial engineering, media arts and animation, and marketing Advanced degree holders including M.S. (44%), M.B.A. (12%), and Ph.D. (25%) Affiliates with extensive industry experience including retired engineers from Boeing, Ball Aerospace, Rockwell, Marquardt, Northrop Grumman, etc. 6

7 Engineering Focus Areas Space Launch Systems Human Space Exploration Robotic Spacecraft and Small Satellites Hypersonic Atmospheric Flight Emerging Commercial Space Markets Revolutionary Aerospace Technologies 7

8 Key Customers and Partners 8

9 Study Overview 9

10 Overview Key Question What is the feasibility of an all-chemical, common-element solution for a human exploration architecture of the Moon, Near Earth Asteroids (NEAs), and Mars using proven, near-term, low-risk technologies? Three Phases 1. Determine the driving design parameters for a Cryogenic Propulsion Stage (CPS) 2. Find strategies to apply these insights to CPS design 3. Generate detailed designs for CPSs and implement them into a comprehensive architecture solution 10

11 Assumptions Use only proven, near-term, low-risk technologies Existing or near-term liquid rocket propulsion systems Passive thermal management of cryogenic propellants (no active systems) Non-propulsive elements were not analyzed in detail Masses selected from literature based on current publically available data Mission operations beyond propulsive maneuvers not considered Treat mission initial mass as only figure of merit All missions start fully fueled in space; launch vehicles and manifests not considered Cost, operations, reliability not analyzed 11

12 Payloads Crew + Suits Number: 4 Mass: 0.60 t (total) Multi-Purpose Crew Vehicle (MPCV) Mass: t Crew Capacity : 4 Duration: 21 days Lunar Lander Mass: 30.0 t Crew Capacity: 4 Duration: 14 days Space Exploration Vehicle (SEV) Mass: 6.5 t Crew Capacity : 2 Duration: 28 days Deep Space Habitat 1 (DSH-1) Mass : 23.5 t Crew Capacity : 4 Duration: 180 days Deep Space Habitat 2 (DSH-2) Mass : 30.0 t Crew Capacity : 4 Duration: 1,000 days 12

13 Missions Outbound TOF = days Outbound TOF = days Outbound TOF = 3-4 days Stay Time = 7 days Stay Time = 30 days Stay Time = days LEO E-M L1 LLO LMO Return TOF = 3-4 days Return TOF = days Return TOF = days Moon Mission NEA Mission Mars Mission 13

14 Parametric Sensitivity Analysis 14

15 Design Variables Design Variable Value Justification 0.75 Low-end CPS mass fraction Propellant Mass Fraction 0.80 Conservative mass fraction 0.85 Ares V EDS-like mass fraction 0.90 Centaur-like mass fraction 0.95 High-end aggressive mass fraction Engine Specific Impulse 448 sec J-2X 451 sec RL10-A sec RL10-B2, MB-60, or NGE Propellant Boil-Off Rate 0.1%/day 0.05%/day 0.01%/day 0.001%/day Centaur boil-off rate achievable via modifications Reasonable near-term boil-off rate with passive system Aggressive boil-off rate with passive system Requires active system 15

16 Concepts of Operations Moon Mission NEA Mission Mars Mission LLO LMO SEV DSH-1 DSH-2 Lunar Lander MPCV CPS 1 MPCV CPS 2a MPCV CPS 2b CPS 1a CPS 1b Direct Direct LEO or L1 Entry LEO or L1 Entry LEO or L1 Direct Entry 16

17 Lunar Mission Strong sensitivity to PMF Little sensitivity to Boil-Off Rate Weak sensitivity to Isp 17

18 NEA Mission Strong sensitivity to PMF Mild sensitivity to Boil-Off Rate Weak sensitivity to Isp 18

19 Mars Mission (orbit only) Strong sensitivity to PMF Strong sensitivity to Boil-Off Rate Weak sensitivity to Isp 19

the propellants lost to boil-off during loiter time?

20 Refueling before Departure What is the sensitivity of initial mass to refueling the stages in orbit before departure in order to replenish the propellants lost to boil-off during loiter time? Initial mass increases by 10% for NEA Mission Initial mass increases by 60% for Mars Mission 20

21 Application to CPS Design 21

22 Goals Investigate options to improve propellant mass fraction. These technologies will significantly benefit all future human exploration mission categories Investigate high performance, near-term in-space chemical rocket engines utilizing LOX and LH2 propellants Investigate options that reduce propellant boil-off rates to minimize propellant losses. This is particularly important for Mars missions Assume a refuelable CPS in LEO. This will mitigate the impact of propellant losses for missions with long wait times before Earth departure 22

23 Structures Design Bypass launch loads by housing CPS inside a fairing Launch CPS partially full (minimal LOX) to reduce sizing structures loads Choice of tank design (e.g. monocoque, common bulkhead stainless steel construction vs. separated aluminum isogrid tanks) can have a factor of 2 impact on tank mass Oversizing propulsion for CPS drives unnecessary loads through the stage, increasing tank mass 23

Subsystems Design Pumps Vaporizers IVF Controller GOX/GH2 Internal Combustion Engine Battery Axial Thrusters Yaw Thrusters Pitch Thrusters Accumulator

control Tank pressurization Power generation Eliminates the need for separate: Attitude control reactants Pressurant tanks and gases Electric power

24 Subsystems Design Pumps Vaporizers IVF Controller GOX/GH2 Internal Combustion Engine Battery Axial Thrusters Yaw Thrusters Pitch Thrusters Accumulator Tanks Integrated Vehicle Fluids (IVF) currently being developed by ULA Boil-off gases from LH2 and LOX tanks collected and stored to provide: Attitude control Tank pressurization Power generation Eliminates the need for separate: Attitude control reactants Pressurant tanks and gases Electric power generation consumables Zegler, F., An Integrated Vehicle Propulsion and Power System for Long Duration Cryogenic Spaceflight, United Launch Alliance,

Tank Design Innovative tank design for passive thermal control can significantly reduce boil-off losses with only modest mass penalties Low-conductivity adapters and tank Settled pressure control Low

25 Tank Design Innovative tank design for passive thermal control can significantly reduce boil-off losses with only modest mass penalties Low-conductivity adapters and tank Settled pressure control Low surface area and thermal mass tank Vapor cooled interface and retention device Vacuum insulation panels Thermodynamic vent system Internal vacuum feed line Minimum penetrations Integrated shield with multi-layer insulation micrometeoroid protection, and vapor cooling Propellant positional management system Sump design and dry feed lines Kutter, B., et Al., Atlas Centaur Extensibility to Long-Duration In-Space Applications, AIAA

maintained at different angles to provide an open path for thermal energy to be directed out into deep space Goff, J., Kutter, B., Zegler, F.

26 Sun Shield ULA is developing a lightweight, deployable sun shield for long term missions Shield is composed of multiple layers to form a concentric cone around the CPS and shadows the tank from both solar and planetary heating Multiple, concentric conical shield layers maintained at different angles to provide an open path for thermal energy to be directed out into deep space Goff, J., Kutter, B., Zegler, F., Bienhoff, D., Chandler, F., Marchetta, J., Realistic Near-Term Propellant Depots: Implementation of a Critical Spacefaring Capability, AIAA,

2 ft MB-60 Engine Specifications Propellants:

engine design to utilize modern design and

Restart-able, high performance, in-space

27 Propulsion Next Generation Engine (NGE) Specifications Propellants: LOX/LH2 Tvac: 30,000 lbf Isp,vac: 465 sec O/F: 5.8 T/W (vac): 60 Area Ratio: 300 Length: 9.2 ft MB-60 Engine Specifications Propellants: LOX/LH2 Tvac: 60,000 lbf Isp,vac: 465 sec O/F: 5.8 T/W (vac): 46 Area Ratio: 300 Length: 10.8 ft Description Notional cryogenic upper-stage engine design to utilize modern design and manufacturing to minimize life cycle cost Likely suppliers are Aerojet, Pratt & Whitney Rockeydyne, XCOR Aerospace Description Restart-able, high performance, in-space expander cycle LOX/LH2 engine A joint development between Pratt & Whitney Rocketdyne and Mitsubishi Heavy Industries, Ltd. 27

28 Architecture Solution 28

29 Architecture Model SpaceWorks created an implementation of its StageSizer tool to parametrically size the CPSs StageSizer is a conceptual level sizing and analysis tool for in-space stages that uses historical mass estimating relationships, physics-based equations, and empirical data Outputs include a multi-level mass breakdown statement that details the contributions of subsystems, margin, propellants, etc., to total stage mass The model also outputs the required vehicle volumes and dimensions StageSizer is integrated into a Mission Scenario and Analysis Tool (MSAT) model The MSAT model determines the required propellant mass for each mission from the required mission ΔVs and boil-off losses An optimizer is used to minimize the total mass of each CPS design while achieving the required payload mass for each mission The optimizer can offload propellant from different stages for each mission to improve payload performance 29

bulkhead tanks Integrated Vehicle Fluids Equipment shelf LH2 tank LOX

30 CPS Alpha Assumptions (Optional) Deployable Sun Shield Multi-Layer Insulation (MLI) Stainless steel pressure-stabilized tanks Common bulkhead tanks Integrated Vehicle Fluids Equipment shelf LH2 tank LOX tank (with LH2 passthrough) Avionics Aft thrust structure 4x NGE Engines 30

CPS Beta Assumptions (Optional) Deployable

Integrated Vehicle Fluids Equipment shelf

31 CPS Beta Assumptions (Optional) Deployable Sun Shield Stainless steel pressure-stabilized tanks Multi-Layer Insulation (MLI) Intertank structure Integrated Vehicle Fluids Equipment shelf LH2 tank LOX tank Avionics Aft thrust structure 4 x MB-60 Engines 31

tanks Integrated Vehicle Fluids Equipment shelf LH2 tank

32 CPS Gamma Assumptions Multi-Layer Insulation (MLI) Stainless steel pressure-stabilized tanks Common bulkhead tanks Integrated Vehicle Fluids Equipment shelf LH2 tank LOX tank (with LH2 passthrough) Aft thrust structure 5 x MB-60 Engines 32

CPS Family CPS Alpha Block 1 CPS Alpha Block 2 CPS Beta Block 1 CPS Beta Block 2 CPS Gamma (Optional) Max Diameter 5.0 m 5.0 m 6.5 m 6.5 m 8.0 m Total Length 19.2 m 21.2 m 28.5 m 31.0 m 32.

33 CPS Family CPS Alpha Block 1 CPS Alpha Block 2 CPS Beta Block 1 CPS Beta Block 2 CPS Gamma (Optional) Max Diameter 5.0 m 5.0 m 6.5 m 6.5 m 8.0 m Total Length 19.2 m 21.2 m 28.5 m 31.0 m 32.1 m Propellant Mass 97.1 t 97.1 t t t t Wet Mass t t t t t Total PMF 92.9% 91.7% 92.8% 91.8% 94.4% Engines 4x NGE 4x NGE 4x MB-60 4x MB-60 5x MB-60 Avg. Boil-Off Rate 0.030%/day 0.015%/day 0.030%/day 0.015%/day 0.030%/day 33

34 Moon Mission Initial Mass = 157 t Lunar Lander MPCV CPS Alpha B1 34

35 NEA Mission SEV Initial Mass = 345 t DSH-1 Initial Mass = 355 t SEV MPCV DSH-1 SEV CPS Alpha B2 MPCV Initial Mass = 180 t DSH-1 MPCV CPS Alpha B2 CPS Alpha B1 CPS Alpha B2 CPS Beta B1 CPS Alpha B1 CPS Alpha B1 35

36 Mars Mission (orbit only) Initial Mass = 660 t DSH-2 MPCV Initial Mass = 690 t CPS Beta B2 DSH-2 MPCV Initial Mass = 345 t DSH-2 MPCV CPS Beta B1 CPS Beta B2 CPS Beta B2 CPS Beta B1 CPS Gamma CPS Beta B1 36

37 Summary and Conclusions 37

38 Design Lessons from Sensitivity Analysis Propellant Mass Fraction is the largest design driver based on the range of values chosen for the design variables for all missions Specific Impulse has a small impact over the ranges available to LOX/LH2 engines Boil-off rate on the long duration CPS in the Mars mission has a large impact on mission masses If there are significant loiter times in LEO or L1 before mission departure, refueling the CPS prior to departure can greatly reduce system mass 38

39 Application to CPS Design Primary structures mass can be reduced by launching the CPS empty as a nonloadbearing payload within a fairing Subsystems masses can be reduced with Integrated Vehicle Fluids Boil-off rates can be reduced passively through several means: Tank design to minimize thermal losses Integrated thermal shielding and vapor cooling Lightweight deployable sun shield for long duration missions High performance near-term propulsion options are available for LOX/LH2 vehicles 39

$Insights from Architecture Development High propellant mass fractions and low boil-off rates are required to enable$ great benefit to development cost, but not without some inefficiencies associated with a highly modular solution

great benefit to development cost, but not without some inefficiencies associated with a highly modular solution

40 Insights from Architecture Development High propellant mass fractions and low boil-off rates are required to enable an allchemical solution but are easily achievable with near-term technologies Using common elements can provide a great benefit to development cost, but not without some inefficiencies associated with a highly modular solution Departing from Earth-Moon L1 can reduce initial system mass and may prove advantageous, but comes with its own challenges 40

41 Conclusions An all-chemical, common-element solution is possible for human exploration missions to the Moon, NEAs, and Mars using proven, near-term, low-risk technologies. Propulsive stages show the highest growth sensitivity to propellant mass fraction. Technologies that reduce dry mass should be prioritized, as these technologies will benefit all future human exploration missions. Existing or near-term technologies mitigate the cost and risk associated with technology development. By using passive thermal control, reliability issues associated with active thermal systems can be avoided. On-orbit fueling can improve mission flexibility. When launched partially full, the CPSs can be supported by a variety of launch vehicles. 41

SPACEWORKS ENTERPRISES, INC. (SEI) www.sei.aero info@sei.

1+770-379-8000 WASHINGTON: 1701 K Street NW, Suite 750 Washington, DC 20006

42 SPACEWORKS ENTERPRISES, INC. (SEI) ATLANTA: 1040 Crown Pointe Parkway, Suite 950 Atlanta, GA USA WASHINGTON: 1701 K Street NW, Suite 750 Washington, DC USA HUNTSVILLE: 1525 Perimeter Parkway NW, Suite 150 Huntsville, AL USA

43 Appendices 43

44 Appendix: Trajectory Analysis 44

45 Missions with Assumptions Outbound TOF = days Outbound TOF = days ΔV = 4,400 m/s Outbound TOF = 3-4 days ΔV = 3,350 m/s ΔV = 3,150 m/s ΔV = 950 m/s ΔV = 2,000 m/s ΔV = 2,200 m/s ΔV = 1,770 m/s ΔV = 560 m/s Stay Time = 7 days E-M L1 185 km LEO 100 km LLO Stay Time = 30 days Stay Time = days 400 km LMO ΔV = 1,050 m/s ΔV = 2,150 m/s ΔV = 2,550 m/s Moon Mission Return TOF = 3-4 days Return TOF = days Return TOF = days NEA Mission Mars Mission 45

46 NEA Candidate Selection SpaceWorks identified 27 candidate NEOs for human missions that have been considered in previous studies SpaceWorks narrowed down this list based on launch date and total required delta-v to 7 candidates for further study SpaceWorks collected orbital parameters and ephemeris data for these 7 NEOs from the JPL HORIZONS online database to analyze missions to these asteroids Candidate NEO Launch Date Earth Departure ΔV (m/s) Post Escape ΔV (m/s) Total ΔV (m/s) Mission Duration (days) 2006 FH36 2/16/2016 3,849 4,165 8, HU4 4/5/2016 3,280 1,980 5, JN1 11/26/2016 3,232 5,260 8, VG 7/21/2017 3,360 2,290 5, CQ36 5/6/2018 3,244 6,213 9, GP2 11/6/2019 3,360 1,570 4, EA9 11/28/2019 3,420 2,120 5, KY26 5/21/2020 3,362 4,164 7, UN12 5/22/2020 3,300 1,450 4, QJ142 4/24/2024 3,490 3,400 6, AO10 9/19/2025 3,320 3,740 7, LN6 12/21/2025 3,330 3,690 7, SG344 4/27/2028 3,340 3,220 6, UQ216 8/15/2028 3,710 3,550 7, Apophis 4/13/2029 3,448 6,601 10, DQ14 8/27/2030 3,770 2,100 5, CG9 8/18/2033 3,530 3,080 6, LG6 1/2/2036 3,270 3,210 6, FR85 9/24/2039 3,610 1,790 5, LC 12/13/2039 3,310 3,310 6, VX25 6/12/2040 3,360 3,870 7, YM9 12/27/2044 4,060 2,360 6, UB17 4/22/2045 3,290 3,450 6, WB 6/1/2050 3,450 3,250 6, QQ56 3/7/2051 3,450 2,470 5, HC 4/20/2054 3,950 3,190 7, JD 11/4/2054 5,320 1,650 6,

47 NEA Mission Trajectory Analysis SpaceWorks used Bullseye, their commercially available interplanetary trajectory tool, to independently check the ΔVs found in literature based on the latest ephemeris data This ephemeris used in Bullseye to generate these results was the latest available from JPL and may be different than the ephemeris used in the literature references ΔV From Literature (m/s) SpaceWorks Calculated ΔV (m/s) Candidate NEO Launch Date Earth Departure Post Escape Total Earth Departure NEO Arrival NEO Departure Total 1998 KY26 5/21/2020 3,362 4,164 7, UN12 5/22/2020 3,300 1,450 4,750 3, , QJ142 4/24/2024 3,490 3,400 6,890 3,350 1,220 1,990 6, AO10 9/19/2025 3,320 3,740 7,060 3,230 2,020 2,110 7, LN6 12/21/2025 3,330 3,690 7,020 3,230 2,230 2,210 7, SG344 4/27/2028 3,340 3,220 6,560 3, ,150 6, UQ216 8/15/2028 3,710 3,550 7,260 3,350 1,800 1,790 6,940 47

48 Mars Mission Opportunities SpaceWorks used Bullseye, their commercially available interplanetary trajectory tool, to perform a sparse sweep of conjunction class Earth-to-Mars trajectories in the timeframe and identify 4 mission opportunities for drill-down investigation. Opportunity Start Date End Date 1 1/1/ /6/ /11/ /16/ /19/2024 1/14/2025 Opportunity 1 Opportunity 2 Opportunity 3 Opportunity 4 4 6/28/2026 2/13/2027 * Subject to maximum time-of-flight constraint of 365 days Within these opportunities, SpaceWorks then determined appropriate launch windows and performed a detailed analysis of each window using Bullseye 48

49 Mars Mission Opportunity km/s 2.2 km/s 2.7 km/s 30 day window 30 day window Delta-Vs required to achieve 30 day launch windows for outbound and return legs: Maneuver ΔV or Velocity Opening Date Closing Date Earth Departure (ΔV) 4,000 m/s 7/9/2020 8/8/2020 Mars Arrival (ΔV) 2,200 m/s 1/30/2021 3/6/2021 Mars Departure (ΔV) 2,550 m/s 7/1/2022 8/5/2022 Earth Entry (maximum velocity) 12.2 km/s 4/12/2023 5/12/2023 Surface stay duration is 480 days to 550 days 49

50 Mars Mission Opportunity km/s 2.3 km/s 2.5 km/s 30 day window 30 day window Delta-Vs required to achieve 30 day launch windows for outbound and return legs: Maneuver ΔV or Velocity Opening Date Closing Date Earth Departure (ΔV) 4,400 m/s 9/2/ /2/2020 Mars Arrival (ΔV) 2,500 m/s 4/10/2023 5/30/2023 Mars Departure (ΔV) 2,300 m/s 7/7/2024 8/11/2024 Earth Entry (maximum velocity) 11.5 km/s 5/8/2025 5/23/2025 Surface stay duration is 400 days to 490 days 50

51 Mars Mission Opportunity km/s 2.55 km/s 2.15 km/s 30 day window 40 day window Delta-Vs required to achieve 30 day launch windows for outbound and return legs: Maneuver ΔV or Velocity Opening Date Closing Date Earth Departure (ΔV) 4,400 m/s 10/11/ /10/2024 Mars Arrival (ΔV) 2,550 m/s 6/8/2025 9/11/2025 Mars Departure (ΔV) 2,150 m/s 7/15/2026 8/24/2026 Earth Entry (maximum velocity) 12.0 km/s 5/11/2027 6/10/2027 Surface stay duration is 300 days to 440 days 51

52 Mars Mission Opportunity km/s 2.5 km/s 2.05 km/s 30 day window 50 day window Delta-Vs required to achieve 30 day launch windows for outbound and return legs: Maneuver ΔV or Velocity Opening Date Closing Date Earth Departure (ΔV) 4,250 m/s 11/10/ /10/2026 Mars Arrival (ΔV) 2,500 m/s 8/2/2027 8/27/2027 Mars Departure (ΔV) 2,050 m/s 8/6/2028 9/25/2028 Earth Entry (maximum velocity) 12.4 km/s 7/22/2029 8/26/2029 Surface stay duration is 340 days to 420 days 52

53 L1 Departure Options SpaceWorks identified four possible departure paths for deep space missions departing from Earth-Moon L1 into a heliocentric orbit These options were compared at different required Earth escape C3 values to determine the optimal mission path and total departure ΔV for each option using in-house cis-lunar trajectory tools L1 to Earth TOF = 3.3 days Direct Escape Moon Flyby (Lunar Gravity Assist) Perigee Maneuver Moon to Earth TOF = 3.1 days L1 to Moon TOF = 2.7 days L1 Maneuver Perilune Maneuver L1 to Moon TOF = 2.7 days Moon Flyby (Lunar Gravity Assist) + Earth Flyby Earth Flyby 53

54 L1 Departure ΔV Requirements NEO Mission C3 = 3.8 km 2 /s 2 Mars Mission C3 = 28.2 km 2 /s 2 Assumptions Model uses two-body equations and patched conic method to determine position, velocity, and time of flight Heliocentric declination of Earth departure velocity vector is always assumed to match that required to for the particular mission, i.e. the Earth-Moon orientation at L1 departure is optimal NEO Optimal Trajectory: Moon Flyby* Total ΔV: 560 m/s Mars Optimal Trajectory: Moon Flyby + Earth Flyby Total ΔV: 1,770 m/s Heliocentric inclination of the Earth departure velocity vector is not considered. Minor adjustments to the spacecraft velocity vector before the flyby maneuvers should allow for an appropriate range of heliocentric inclinations to be achieved. All maneuvers assumed to be instantaneous changes in velocity with no gravity losses * Departure trajectory without Earth flyby may significantly limit the departure heliocentric inclination available 54

55 ΔV Losses Overview All ΔVs calculated are the ideal, instantaneous ΔV required for each mission maneuver For most mission legs this assumption is sufficient, however for low thrust-to-weight stages leaving Low Earth Orbit, significant gravity losses can be incurred The impact of these losses can be reduced to a function of stage Thrust-to-Weight (T/W) and maneuver Mass Ratio (MR) at a given engine specific impulse and initial LEO Mass ratio = initial mass / final mass for propulsive maneuver Burn time can be expressed solely as a function of T/W and MR For this study, a 400 km by 400 km LEO and 465 sec engine vacuum specific impulse were assumed for all missions starting in LEO A simulation was set up in POST 3D to determine the ΔV losses from LEO based on different combinations of T/W and MR First, for the given MR, the ideal ΔV was calculated Next, the simulation was executed and the final orbital specific energy was determined in POST 3D Based on this orbital energy, the theoretical initial velocity in LEO was determined From the initial velocity, an actual ΔV was calculated This actual ΔV is compared to the ideal ΔV for the given MR 55

56 ΔV Losses Results ΔV losses increase as initial T/W decreases increasing stage thrust will directly reduce gravity losses ΔV losses increases as maneuver MR increases stages with low MR will suffer less gravity losses 56

57 Appendix: CPS Designs 57

58 CPS Alpha Block 1 WBS 5.0 m Category Mass (lbm) Mass (kg) 19.2 m 1.8 m 7.9 m 3.7 m 1.8 m Structure 7,660 3,480 Propulsion 2,730 1,240 Thermal Control Avionics Miscellaneous Margin (30%) 3,790 1,720 Dry Mass 16,430 7,450 Residuals, Reserves, and Buildup 3,150 1,430 Consumables 0* 0* RCS Propellants 0* 0* Inert Mass 19,580 8,880 Usable Fuel 30,900 14,020 Usable Oxidizer 178,930 81,160 Startup Losses 1, Wet Mass 230, ,530 * IVF allows use of main propellants for pressurization and attitude control Category Value Total Propellant Mass 214,020 lbm 97,080 kg Usable Propellant Mass Fraction 91.0% Total Propellant Mass Fraction 92.9% Average Boil-Off Rate 0.03%/day 58

59 CPS Alpha Block 2 WBS 5.0 m Category Mass (lbm) Mass (kg) 21.2 m 2.0 m 1.8 m 7.9 m 3.7 m 1.8 m Structure 7,660 3,480 Propulsion 2,730 1,240 Thermal Control 3,190 1,450 Avionics Miscellaneous Margin (30%) 4,450 2,020 Dry Mass 19,290 8,750 Residuals, Reserves, and Buildup 3,150 1,430 Consumables 0* 0* RCS Propellants 0* 0* Inert Mass 22,440 10,180 Usable Fuel 30,900 14,020 Usable Oxidizer 178,930 81,160 Startup Losses 1, Wet Mass 233, ,830 * IVF allows use of main propellants for pressurization and attitude control Category Value Total Propellant Mass 214,020 lbm 97,080 kg Usable Propellant Mass Fraction 89.9% Total Propellant Mass Fraction 91.7% Average Boil-Off Rate 0.015%/day 59

60 CPS Beta Block 1 WBS 28.5 m 6.5 m 5.0 m 2.3 m 9.9 m 2.3 m 5.6 m Category Mass (lbm) Mass (kg) Structure 17,500 7,940 Propulsion 6,490 2,940 Thermal Control 1, Avionics Miscellaneous Margin (30%) 8,140 3,700 Dry Mass 35,300 16,010 Residuals, Reserves, and Buildup 6,720 3,050 Consumables 0* 0* RCS Propellants 0* 0* Inert Mass 42,020 19,060 Usable Fuel 65,860 29,870 Usable Oxidizer 381, ,270 Startup Losses 2,240 1,020 Wet Mass 492, ,220 * IVF allows use of main propellants for pressurization and attitude control Category Value Total Propellant Mass 456,810 lbm 207,205 kg Usable Propellant Mass Fraction 91.0% Total Propellant Mass Fraction 92.8% Average Boil-Off Rate 0.03%/day 60

61 CPS Beta Block 2 WBS 31.0 m 6.5 m 5.0 m 2.5 m 9.9 m 5.6 m Category Mass (lbm) Mass (kg) Structure 17,500 7,940 Propulsion 6,490 2,940 Thermal Control 6,310 2,860 Avionics Miscellaneous Margin (30%) 9,465 4,290 Dry Mass 41,020 18,610 Residuals, Reserves, and Buildup 6,720 3,050 Consumables 0* 0* RCS Propellants 0* 0* Inert Mass 47,740 21,650 Usable Fuel 65,860 29,870 Usable Oxidizer 381, ,270 Startup Losses 2,240 1,020 Wet Mass 497, ,810 * IVF allows use of main propellants for pressurization and attitude control Category Value Total Propellant Mass 456,810 lbm 207,205 kg Usable Propellant Mass Fraction 90.0% Total Propellant Mass Fraction 91.8% Average Boil-Off Rate 0.015%/day 61

62 CPS Gamma WBS 8.0 m Category Mass (lbm) Mass (kg) 32.1 m 2.8 m 14.9 m 6.7 m 2.8 m Structure 32,370 14,680 Propulsion 8,257 3,740 Thermal Control 2,800 1,270 Avionics 1, Miscellaneous 1, Margin (30%) 13,780 6,250 Dry Mass 59,710 27,080 Residuals, Reserves, and Buildup 14,730 6,680 Consumables 0* 0* RCS Propellants 0* 0* Inert Mass 74,440 33,760 Usable Fuel 144,410 65,500 Usable Oxidizer 837, ,920 Startup Losses 4,910 2,230 Wet Mass 1,061, ,420 * IVF allows use of main propellants for pressurization and attitude control Category Value Total Propellant Mass 1,001,630 lbm 454,330 kg Usable Propellant Mass Fraction 92.5% Total Propellant Mass Fraction 94.4% Average Boil-Off Rate 0.03%/day 62

63 Appendix: Architecture Results 63

64 Moon Mission Parameter Value Payload Mass MPCV Lunar Lander Crew (4) + Suits Payload Adapter CPS Version Fuel Offload Oxidizer Offload CPS Propellant Mass at Ignition CPS Wet Mass at Ignition Total Mass at Ignition 52,080 kg 21,250 kg 30,000 kg 600 kg 230 kg Alpha B1 0 kg 0 kg 97,080 kg 104,530 kg 156,610 kg Initial Thrust-to-Weight at Ignition 0.35 Gravity Losses 1.5% Total Burn Time Videal 13.3 min 4,150 m/s 64

65 Comparison of NEA Mission Options LEO Departure Option 1 LEO Departure Option 2 L1 Departure Number of Stages Payload Mass (kg) 52,080 52,080 52,080 MPCV 21,250 21,250 21,250 Deep Space Habitat (Block 1) 23,500 23,500 23,500 Space Exploration Vehicle 6,500 6,500 6,500 Crew (4) + Suits Payload Adapter Total Initial Mass (kg) 344, , ,220 Stage 1 Stage 2 Stage 3 Stage 1 Stage 2 Stage 1 Stage 2 Element Alpha 1 Alpha 1 Alpha 2 Beta 1 Alpha 2 Alpha 1 Alpha 2 Average Boil-Off Rate (%/day) 0.030% 0.030% 0.015% 0.030% 0.015% 0.030% 0.015% CPS Mass at Ignition (kg) 99,260 99,260 92, ,070 92,980 31,690 92,980 Ignition T/W Total Burn Time (min) Videal (m/s) 1,340 2,040 3,650 3,400 3, ,650 Gravity Losses (%) 1.0% 1.0% < 0.1% 1.5% < 0.1% < 0.1% < 0.1% Fuel Offload (%) 5.8% 5.8% 3.6% 6.1% 3.6% 78.1% 3.6% Oxidizer Offload (%) 5.6% 5.6% 15.5% 6.1% 15.5% 78.1% 15.5% 65

66 Comparison of Mars Mission Options LEO Departure Option 1 LEO Departure Option 2 L1 Departure Number of Stages Payload Mass (kg) 52,080 52,080 52,080 MPCV 21,250 21,250 21,250 Deep Space Habitat (Block 2) 30,000 30,000 30,000 Crew (4) + Suits Payload Adapter Total Initial Mass (kg) 657, , ,490 Stage 1 Stage 2 Stage 3 Stage 1 Stage 2 Stage 1 Stage 2 Element Beta 1 Beta 1 Beta 2 Gamma Beta 2 Beta 1 Beta 2 Average Boil-Off Rate (%/day) 0.030% 0.030% 0.015% 0.030% 0.015% 0.030% 0.015% CPS Mass at Ignition (kg) 233, , , , , , ,910 Ignition T/W Total Burn Time (min) Videal (m/s) 1,650 2,810 4,250 4,620 4,250 1,770 4,250 Gravity Losses (%) 1.0% 1.5% < 0.1% 5.0% < 0.1% < 0.1% < 0.1% Fuel Offload (%) 0.0% 0.0% 9.0% 0.0% 9.0% 44.3% 9.0% Oxidizer Offload (%) 0.0% 0.0% 38.4% 0.0% 38.4% 44.3% 38.4% 66

A STUDY OF CRYOGENIC PROPULSIVE STAGES FOR HUMAN EXPLORATION BEYOND LOW EARTH ORBIT

A STUDY OF CRYOGENIC PROPULSIVE STAGES FOR HUMAN EXPLORATION BEYOND LOW EARTH ORBIT Mark Schaffer SpaceWorks Enterprises, Inc., United States, mark.schaffer@sei.aero Chauncey Wenner United Launch Alliance,