Power, Propulsion, and Thermal Preliminary Design Review James Black Matt Marcus Grant McLaughlin Michelle Sultzman

Transcription

1 Power, Propulsion, and Thermal Preliminary Design Review James Black Matt Marcus Grant McLaughlin Michelle Sultzman

2 Outline 1. Crew Systems Design Selection 2. Thermal Requirements and Design 3. Power Requirements and Design 4. Reaction Control System Design 2

3 Crew Systems Design Selection Selection dictated by placement requirements for thrusters and propellant tanks o o o Able to place thrusters around center of mass plane Crew Systems capsule modified to expand lower cavity between pressure vessel and hull to hold propulsion system New pressure vessel removes unused pressurized space, minimizing impact on crew systems layout Old Pressure Vessel (in green) New pressure vessel (in green) 3

4 Thermal Control System Requirements Must maintain cabin temperature in following cases: o o o o o Full sun (Translunar) Eclipse (LEO/LLO) Lunar surface - Dawn/Dusk/Polar Lunar surface - 45 Sun Angle Lunar surface - Noon Equatorial 4

5 Thermal Control System NASA requirement: Crew cabin temperature within range of 292 K to 300 K Simple heat exchanger between cabin air and fluid that circulates through external radiator o Reduces heat generation from refrigeration cycle o Limits radiator temperature to approximately cabin temperature Assuming all radiative surfaces coated with white paint o α= 0.2 o ε = 0.8 5

6 Multi-Layer Insulation Surface area of craft covered with multi-layer insulation determined by considering case of thermal equilibrium during eclipse T rad = Q int ε σ A tot + T env Q int = 1445 W, T env = 4 K A radiator = A total = 25 m 2 : K A radiator = 6 m 2 : K A radiator = 4.25 m 2 : K 6

7 Multi-Layer Insulation During eclipse: o Designing the craft to have an equilibrium temperature within acceptable range solely through use of insulation left too small of an area to effectively radiate heat during other phases of the mission o Instead craft will have an equilibrium temperature of 270 K, and will use a 500 W space heater to increase temperature of pressure vessel Final design: 19 m 2 of total surface area of 25 m 2 will be covered with MLI 7

8 Radiators Considered Additional radiators were necessary Initially considered radiators deployed like flower petals at the top of the cone o This design did not yield enough surface area for all thermal profiles 8

9 Radiators Selected Three double-sided circular radiator arrays, 8.44 m 2 on each side Radiator arrays, along with solar array of the same size (to be discussed later) will be folded out from the heat shield before docking with orbital propulsion module o Each array spaced 90 apart o Arrays not obstructing ingress/egress Each radiator contains two channels of coolant which can be turned on and off independently, allowing number of radiators used to vary in increments of 0.5 9

10 Radiator Array Configuration Radiator Solar Array Radiator Radiator 10

11 Radiators Radiators can rotate axially in order to have highest temperature control resolution possible Thermal equilibrium calculations were performed assuming radiator panels were oriented perpendicular to lunar surface o Rotating panels slightly will alter illuminated area of craft, and therefore increase or decrease equilibrium temperature for more precise thermal control 11

12 Thermal Equilibrium Calculations Assumed that all radiated heat from spacecraft is radiating directly into space Assumed that panels perpendicular to lunar surface will radiate from both sides o o 65% of heat will radiate into space 35% of heat will radiate toward Lunar surface Illuminated area changes with solar angle Internally generated heat comprised of power requirements of different systems at each phase and 116 W heat load per crew member (based on average 2400 kcal daily diet) 12

13 Spacecraft Thermal Control Cases Case Q int (W)* Illuminated Area (m 2 ) T env (K) Number of active panels T rad (K) Eclipse (space) Full Sun (space) Lunar Surface Dawn/Dusk/ Polar Lunar Surface - 45deg Solar Angle Lunar Equatorial Noon (space) 180 (moon) (space) 215 (moon) (space) 380 (moon) * Values presented on next slide 13

14 Mission Phase Power Requirements RCS (W) Life Support (W) Avionics (W) Lighting (W) Thermal (W) Total (W) LEO checkout Cis-Lunar Space LLO Loiter Lunar D/A Lunar Surface Ops Earth EDL W average, 1200 W Peak 14

15 Power Generation Power Systems Investigated o o o Lithium Ion Battery Powered LOX/LH2 Fuel Cell Solar Panels with rechargeable Lithium Ion Battery o Solar panels not active during launch, entry, eclipse Trade Study performed on total mass of each system for mission duration 15

16 Power Generation System Mass Trade Study 16

17 Power System Sizes Solar Panels/ Li-Ion Battery Solar Array: 8.44 m 2 Battery: 0.01 m 3, 10 kg Apollo Fuel Cell LOX tank: 0.08 m 3 LH2 tank: m 3 Fuel cell: 0.16 m 3

18 Power System Design Choice Solar Panel and Battery System chosen o o Smaller mass than fuel cell Battery large enough to power during phases when solar panels do not produce power 18

19 Propulsion Requirements Translational delta-v of 50 m/s Attitude hold in +/- 5 degree deadband for 3 days for return to earth Overcome 500 Nm pitch and yaw moments during ~3 minute reentry Rotate Spacecraft 10 times 180 degrees in <30 sec 19

20 Coordinate System +X +X +Z +Y 20

21 Moments of Inertia Assuming a hollow cone with wall thickness of 1/3 of the maximum radius o r = m o h = m o m = 4795 kg Moments are CCW about axis Ix = 1/6*m*r 2 = 2140 [kg*m 2 ] Iy = Iz = 1/3*m*(r 2 /4+h 2 ) = [kg*m 2 ] 21

22 Translational Propulsion Rocket Equation : V f -V i = I sp *g*exp(m i /M f ) I sp ~300 sec M i = 4795 kg M f = 4795-(Propellent Mass) Delta v = 50 m/s Account for thrusters firing at 50 degree angle from vertical for +x translation Propellent Mass = 125 kg 22

23 Translational Propulsion Maneuvers +X Front View Side View 23

24 Translational Propulsion Maneuvers -X Front View Side View 24

25 Translational Propulsion Maneuvers +Y, -Y Top View Front View 25

26 Translational Propulsion Maneuvers +Z, -Z Top View Side View 26

27 Rotational Propulsion Maneuvers: Roll roll about +X, -X Top View 27

28 Rotational Propulsion Maneuvers: Yaw Top View Left Side View Yaw about +Z, -Z Front View Right Side View 28

29 Rotational Propulsion Maneuvers: Pitch Top View Left Side View Pitch about +Y, -Y Front View Right Side View 29

30 Thrust Requirements: Pitch, Yaw The placement of the thrusters was determined from the constraint of the largest necessary moment for control: pitch and yaw Maximum aerodynamic moment for yaw and pitch: 500 Nm Assuming some control over the location of the center of gravity (CG), it is best to have the thrusters in the xy-plane at about 3/4 from the top 30

31 Thrust Requirements: Pitch and Yaw In atmospheric entry, translation forces do not need to be balanced Overcome 500 Nm pitch and yaw moments Two thrusters firing in the +x direction Thruster per Draco: 250 N Moment arm: 1.0 m F=mu m = 400 N 300s 9.81 m s 2 = kg/s Fraction of time burning: T f = 500/(800*1) Mass of propellent = T f *180* m *2 = kg 31

32 Thrust Requirements: Roll 1 2 τ I t2 + θ ot = θ θ o Assume angular rate = t=0 and t=30 Need equivalent of -30 Nm for 15 sec and 30 Nm for 15 sec Thrust of N per thruster (for perfectly coupled torque) Performed by two 400N Draco thrusters for 1.65 sec at 100% duty cycle 32

33 Attitude Trajectories Trajectories for various torques with initial angle and angular rate offsets Able to stay within 5 degree deadband with -75 Nm torque Well within limits of attitude control system 33

34 Thrust Requirements: Deadband Pitch and yaw performed by: o One thruster in the -x direction; arm = 1.4 m; thrust = 26.8 N o Two thrusters in the +x direction; arm 1 m; thrust = 18.8 N Voyager 1 and 2 have used only 50 kg of propellent since 1977 to maintain deadband 5 degree half deadband width = rad Fuel consumption decreases with increasing deadband width as shown above Can assume propellent used is very small for 3 days - on the order of 1 kg 34

35 Thruster Selection SpaceX Draco thrusters were selection: Similar mission design requirements Thrusters are flush with outer shell 400 N thrust covers all requirements 300 sec Isp Operates at different duty cycles for each reaction control objective Maneuver Thrust Duty Cycle Draco Thruster duty cycle = thrust time thrust time + drift time Pitch 250 N 63% Yaw 250 N 63% Roll 27 N/ 19 N 6.7%/4.9% Deadband 11 N 2.7% 35

36 Propellent Draco Thrusters: 300 sec Isp, 400 N thrust Nitrogen Textroxide (1.443 g/cm 3 ) ; Monomethylhydrozine (.875 g/cm 3 ) Both liquids 20 C Maneuver Propellent Mass (kg) Oxidizer Mass (kg) Fuel Mass (kg) 50 m/s delta v in x direction Deadband attitude hold for 3 days 500 Nm pitch and yaw moments 180 degree roll in <30 sec; 10 times ~1 kg ~.7 ~ kg Total

37 Fuel and Oxidizer tanks Mass of tank = 299.8*(vol. m 3 ) + 2 Assume 2 mm thickness for tanks Fuel and Oxidizer stored as liquids at 20 C Oxidizer Tank empty Mass (kg) Fuel Tank empty Mass (kg) Oxidizer Tank inner Volume (m 3 ) Fuel Tank inner Volume (m 3 )

38 References Akin, David et Al. Minimum Functionality Lunar Habitation Element. University of Maryland Space Systems Laboratory Print. Akin, David. (2012) Rocket Propulsion [PDF]. Retrieved from Akin, David. (2012) Power Systems Design [PDF]. Retrieved from Akin, David. (2012) Thermal Analysis and Design [PDF]. Retrieved from Allen, Christopher et Al. Guidelines and Capabilities for Designing Human Missions. TM NASA JSC, Print. Dragon Overview. SpaceX. Website < Foley, John. Control of a Spacecraft Using a Reaction Control System. < Gilmore, D.G. Spacecraft Thermal Control Handbook. AIAA

39 References Liquid Rocket Propellants. Website < Solo lunar flyby using standard Falcon 9 and Dragon. Forum post. 18 Dec < Typical Onboard Systems. Basics of Space Flight: Section II. Jet Propulsion Laboratory. Website < Thruster Mass Estimation. Delft University of Technology. Website < Washay, Marvin, Prokopius, Paul R. The Fuel Cell in Space: Yesterday, Today and Tomorrow. TM London: NASA, Print. 39