Preliminary Design Review

Transcription

1 Establishing a Recurring Human Presence on the Moon Preliminary Design Review

2 Overview Preliminary Design Review of systems aboard low-cost lunar lander Thermal System Maintain cabin temperature throughout mission Propulsion System 6 DOF control Power System 13 day energy storage requirement Supports all mission phases

3 Requirements 3 Crew Members 10 day mission (+3 contingency days) 3 days transit 4 days on lunar surface 3 days return to Earth Plan for 13 days total (includes 3 contingency days) 4795 kg total mass for lunar lander

4 Reaction Control System (RCS) Requirements 1. Full 6 DoF control 2. Translational Δv of 50 m/sec 3. Three day attitude hold in dead band 4. Overcome 500Nm aerodynamic entry moment roll rotation in 30 sec on entry Must contour to vehicle to survive atmospheric entry

5 RCS Trades Summary Technology Pro Con Cold Gas Inexpensive Low Energy Density and/or ISP Solid Rocket Motor Simpler, Fewer Parts Cannot Be Turned Off Once Ignited Nuclear High ISP, Thrust Lethal Without Excessive Shielding Plasma Extremely High ISP Insufficient Thrust Liquid Rocket Motor Controllable, Good Performance Complex * No need for a detailed trade study, liquid fueled rocket is the only reasonable option

6 RCS Liquid Thruster Design - Overview Assumed layout (Req. 1) 8 identical thrusters (top) 4 identical thrusters (bottom) Approximate pitch/yaw moment arm of 1.5 m (top) Required maximum force per top thruster is N (Req. 4) Bottom thrusters only used for minimal on orbit translation z z y Approximate Center of Mass x

7 RCS Liquid Thruster Design - Sizing Exit Velocity: Use Monomethylhydrazine (MMH) and Dinitrogen Tetroxide (NTO) Known Variables : R = 8314 J/mol*K γ = 1.2 M = Weighted Average of MMH and NTO with 1.67 mix ratio = 74.8 g/mol V e = 1735 m/s Assumed Variables : P 0 = 3.5 MPa P e = 5 kpa

8 RCS Liquid Thruster Design - Sizing r = e (-Δv/ve) = (Req. 2) m p = 4795*(1-r) = kg m MMH = 51.0 kg, m NTO = 85.2 kg, T t = 3398 K Mass Flow Rate: T = ṁ*v e + P e *A e Expansion Ratio : From Above Eqn s A t = 2.50(10-5 ) m 2 A e = 1.35(10-3 ) m 2 ṁ = kg/s

9 RCS Liquid Thruster Design Fluid Transfer Helium pressurized P He_0 = 31 MPa Fluid Mass (kg) Volume (m 3 ) MMH NTO Helium Mechanical regulation to maintain MMH/NTO tanks at 3.5 Mpa Power required for injector valves and propellant preheating to ensure reliable/controlled combustion Thrust burst requires ~8 W per thruster Thrust hold requires ~5 W per thruster

10 RCS Roll Performance Roll rate = 2* * / I = 16.9 deg per sec Torque = 2*166.7 N*0.25 m = 83.3 Nm Roll angle = 180 Moment of Inertia = ~ 6000 kg*m 2 Time to roll sec Meets Req. 5

11 RCS Operation z y z y z y Roll Pitch Yaw z y z y Y-Translation Z-Translation

12 RCS Future Work Iterate as more detailed estimates become available: Center of Mass Moment of Inertia Revisit properties when level of detail allows for component level design/selection Mass breakdown Power requirements

13 Thermal System Requirements Design thermal control system (with radiator temperatures, sizes, and design locations on vehicle) to maintain cabin temperatures in following cases: Full sun (translunar) Eclipse (Earth/Moon orbit) Lunar surface dawn/dusk/polar Lunar surface 45 sun angle (high latitudes/ midmorning or midafternoon) Lunar surface noon equatorial

14 Thermal Control System Design Drivers * Environmental Temperature variance (excluding vehicle re-entry) Total External Heat Dissipation Requirements (varied) Total Internal Heat Dissipation Requirements (6.1 kw) Crew and Equipment Temperature Requirements (~298 K) Spacecraft architecture (radiation surfaces) *Adapted and Modified from [1]

15 Internal Heat Generation Assumptions All power consumed is converted to heat Each crew member outputs 100 W Investigate both possible continuous and stilted increase scenarios from previous measures (i.e. open module during EVAs) Continuous- Must remove 6100 W continuously or 7188 MJ over course of mission and contingency days Stilted - Must be able to remove 708 MJ during initial lunar travel and 927 MJ for return travel + 3 days contingency 28 MJ energy remaining margin reserve (required to vaporize cabin atmosphere LOX and LN 2 )

16 External Conditions Assume heat intercept during Inter-body travel at 1 AU (distance from earth to moon AU) Assume Solar Flux reflected from moon during inter-body travel minimal

17 Maintaining Cabin Temperature White Paint Exterior Coating Alone - Unsuitable AZ Low Alpha White α = 0.09 ± 0.02, ε = 0.91 ± 0.02 [2] Lunar Cases [3] Solar Angle ( ) Lunar Surface (K) Temperature Equilibrium (K) Local Midnight N/A Polar Outpost Day Typical Mid-Latitude Equatorial Noon Inter-body Travel (Between Earth and Moon) No Sun (Eclipse) 235 With Half of Craft Illuminated 258 With Bottom of Craft Illuminated 248 With Top of Craft Illuminated 267 Whole Illumination via earth shine 330

18 Interior Heat Removal Analysis Cryogenic Thermal Storage (Solid H 2 O) Unfeasible, Phase Change Material Alone requires over 1600 kg to be regenerated Daily Heat Exchange System utilizing salt solution Potential energy-providing source Mechanical interior parts required to prevent scaling (failure issue) Mass issues Ammonia Heat Exchanger Highest suitable heat capacity(4.7 j/g K) for liquid (easiest to transport) Extensive health concerns if leaks into cockpit Potential combustion issues if exposed and catalyzed by metal

19 Counter-Flow N-Butanol & H 2 O HX N-Butanol less harmful, but reduction in heat capacity (2.299 J/g K) Area requirements too large when attempting to dissipate heat at sufficient rate Heat to remove 6100 W Inner Loop H2O T in 297 K T out 296 K Mass Flow g/s Outer Loop N-Butanol T in 156 K T out 224 K Mass Flow 7 g/s Area req. for HX m 3

20 Heat Pipes Low Mass requirements (0.25 kg/m) [5] High Wattage Removal Flexibility Issues Capillary Action only suitable in low gravity situations Current Loop Heat Pipe System suitable, total length 1.6 m loop system and 3mm dia. piping with 100W per system removal [6] 61 independent systems and a total length of 97.6 m (interior shell Surface Area requirement of 5.49 m 2 ) 24.4 kg total mass No moving parts, highly reliable Adapted from [6]

21 Lunar Radiating Panels Accordion Panels allow suitable Spacecraft Temperatures during Lunar Sorties Lunar Cases [3] Solar Angle ( ) Lunar Surface (K) Temperature Eq. (K) Local Midnight N/A Polar Outpost Day Typical Mid-Latitude Equatorial Noon Number of Panels 10 Panel Surface Area 11.2

22 Possible Future Energy-Reducing Scenarios Utilize heat dissipated from Avionics Systems to warm-up cockpit before suit removal Utilize fuel-cell-produced water in heat exchanger as secondary thermal control system

23 Power System and Energy Storage Trade Studies Various types of batteries Batteries vs. fuel cells Fuel cells vs. solar array with batteries Exact power requirements are unknown, as lights, avionics, computers, comms have not yet been designed Trade studies examine a range of possible power requirements

24 Various Battery Types kg Specific Energy kwhr Lead Acid Alkaline NiMH NiCad Lithium Ion Lithium Thionyl Chloride

25 Various Battery Types Energy Density 60 Lead Acid m kwhr Alkaline NiMH NiCad Lithium Ion Lithium Thionyl Chloride

26 Battery Trade Study Results Over a range of energy storage requirements, two battery types minimize mass and volume: Lithium Thionyl Chloride Lithium Ion Lithium Thionyl Chloride is only available in small sizes Lithium Ion is the best choice for battery type But even with the most energy dense battery, the mass and volume are too large

27 Battery vs. Fuel Cell Specific Energy kg Lithium Ion Fuel Cell kwhr

28 Battery vs. Fuel Cell m 3 Energy Density kwhr Lithium Ion Fuel Cell

29 Fuel Cell vs. Solar Array Type of solar array used in trade study: Concentrator Triple-Junction Compound Solar Cell Manufactured by Sharp Corporation Used on GOSAT 40% conversion efficiency; 656 W/kg Assumption: Solar array exposed to 10 days of continuous sun Requirement: 3 contingency days worth of Lithium Ion battery energy storage During an appropriately planned contingency scenario, the solar array should be assumed to be inoperative

30 Fuel Cell vs. Solar Array Specific Energy kg kwhr 3 days Lithium Ion + Solar Cells Fuel Cell

31 Fuel Cell vs. Solar Array Battery mass at 3 days energy storage equals fuel cell mass at 13 days energy at 270 kwhr 270 kwhr / 72 hr = 3.75 kw average (may not be enough for required power in a contingency scenario) Due to 3 contingency days energy storage requirement, fuel cells are preferred to a battery/solar array combination

32 Power System Design Until detailed analysis of avionics and computer systems are performed, assume an average power requirement of 6 kw 6 kw * 13 days = 1872 kwhr energy storage 28 VDC output Byproduct of fuel cells is water. Water can be consumed by the crew which decreases amount of water to be carried in the water system at start of mission.

33 Power System Design Mass (kg) Volume (m 3 ) LOX LH LOX tank LH2 tank Reactor Total Reactor dimensions: 0.36 x 0.38 x 1.0 m 3

34 Overall Design - Exterior Front Rear Windows (3x) 25 EVA Hatch

35 Overall Design Interior Parachute Landing Controls / Avionics Food Storage N 2 Tank Reaction Control System Waste Management Air Filter / Dehumidifier Fuel Cells (not visible) O 2 Tank Water Storage

36 Overall Design - Landing Exterior Height: 3.98 m Interior Height: 2.37 m Sight Lines (3 windows evenly spaced around SC) Ingress / Egress Hatch Diameter: 1.0 m 41.9 Exterior Diameter: 3.57 m Interior Diameter: 3.13 m Lunar Surface

37 Mass Budget Component Mass (kg) Crew Systems 1323 Energy storage system (fuel cells) 1130 Reaction Control System 188 Thermal controls 24 Total kg remaining for additional systems and structure

38 Conclusions systems meet requirements for mission Design analysis of remaining systems is needed to determine more precise power requirements and moment of inertia Systems can then be integrated to reduce mass, volume and power requirements Fuel cells can supply thermal and water systems

39 References [1] M.N. De Parolis & W. Pinter-Krainer. Current and Future Techniques for Spacecraft Thermal Control 1. Design drivers and current technologies Thermal Control and Heat Rejection Section, ESTEC, Noordwijk, The Netherlands [2]AZ Technology SPECIALISTS IN MATERIALS AND APPLICATIONS.Spacecraft Thermal Control and Conductive Paints/Coatings* and Services Catalog. Available Online. [3] D. Akin. Thermal Analysis and Design.ENAE 483/788D - Principles of Space Systems Design. University of Maryland, [4]D. Gilmore, M. Donabedian.Spacecraft Thermal Control Handbook: Volume II: Cryogenics. American Institute of Aeronautics & Astronautics. Jan [5] D. Reay, P. Kew. Heat Pipes: Theory, Design and Applications.5 th ed.butterworth-heinmann. Boston, MA [6] Riehl, R. R., Siqueira, T. C. P. A., Heat transport capability and compensation chamber influence in loop heat pipes performance, Applied Thermal Engineering, Elsevier, Science Press, Vol 26/11-12, pp , ISSN , [7] J. A. Stark, K. E. Leonhard, F.O. Bennett. CRYOGENIC THERMAL CONTROL TECHNOLOGY SUMMARIES. NASA CR Dec 1974

40 References [8] Calhoun, Philip C. Entry Vehicle Control System Design for the Mars Smart Lander. AIAA Atmospheric Flight Mechanics Conference and Exhibit. Monterey, California 5-8 August [9] Crisp, R. and Keene, D. Apollo Command and Service Module Reaction Control By the digital Autopilot. MIT Instrumentation Laboratory. May [10] Dyakonov, Artem A. Aerodynamic Interference Due to MSL Reaction Control System. AIAA Thermophysics Conference. San Antonio, Texas, June [11]

41 References [12] Fuel Cell Handbook, 7 th Ed. US DoE. National Energy Technology Laboratory. November, [13] B. McKissock, P. Loyselle, E. Vogel. Guidelines on Lithium-Ion Battery Use in Space Applications. NASA/TM May, [14] Space Shuttle Fuel Cell Power Plants wrplants.html [15] N. Fatemi, et. al., Solar Array Trades Between Very High- Efficiency Multi-Junction and Si Space Solar Cells 28 th IEEE PVSC, Anchorage, Alaska, September, [16] Sharp Develops Concentrator Solar Cell with World s Highest Conversion Efficiency of 43.5% Press release. May, 2012.