Power, Propulsion and Thermal Design Project. Jesse Cummings Shimon Gewirtz Siddharth Parachuru Dennis Sanchez Alexander Slafkosky

Transcription

1 Power, Propulsion and Thermal Design Project Jesse Cummings Shimon Gewirtz Siddharth Parachuru Dennis Sanchez Alexander Slafkosky

2 Mission Itinerary Days 1-3: Voyage to moon Days 4-7: On the lunar surface o + 3 Contingency Days Days 8-10: Voyage back to Earth

3 Requirements Gross mass 4795 kg Must support all mission phases: LEO checkout, Cis-lunar space, LLO loiter, Lunar descent and ascent, Lunar surface operations, Earth EDL. Must be capable of limited 6 DOF control Must maintain cabin temperatures in Full sun, Eclipse, Lunar surface dawn/dusk/polar, Lunar surface 45 sun angle, Lunar surface noon equatorial

4 Crew Systems Capsule Design Selection We approached the capsule selection with a focus on minimizing power requirements and gross mass. With this approach we chose the design which was kg and which used 116 Watts per day. These power and mass requirements are 69% and 9% lower than our second lowest values.

5 Crew Metabolic Heat We assumed that each crew-day there is a total of 348 Watts of metabolic heat radiated per day based on the values from the ECLIPSE slides in the thermal lecture, assuming three crew members.

6 Power: Requirements 1. Account for two phases of mission a. In sun-light b. In darkness 2. Possible Scenarios a. 13 days of darkness b. 13 days of light c. Nominal Case i. 4 Days Dark on Moon ii. 7 days Dark on Moon 3. Examine available technologies 4. Select combination to ensure power to system through energy storage or power generation

7 Power: Assumptions and Reasoning Assumptions 1. Insolation Constant=1394 W/m^2 is actually constant over the range of the mission 2. Technologies examined can perform to capabilities W of additional power assumed for thruster solenoids as a conservative estimate. Reasoning 1. Examine worst case to determine best energy storage device 2. Examine best case to determine best power generation device 3. Create a combination to meet requirements

8 Power: Selections I Battery and Cells 1. Ni-Cd 2. IPV 3. CPV 4. Ni-MH 5. Li-Ion 6. NaS Solar Panels 1. GaAs 2. 2 Junction 3. 3 Junction 4. 4 Junction 5. Single Crystal Si 6. CIGS

9 Power: Selections I These Technologies were excluded: Type - Reason Nuclear and Solar Thermal - Size and Mass Flywheel - Low Watt Hours per SOA H2-O2 Fuel Cell - Storage of H2 and O2 Chemical Thermal - Scalability

10 Power: Trade Study on Mass of Energy Cells

11 Power: Trade Study on Mass of Batteries

12 Power: Trade Study on Volume of Battery and Cells

13 Power: Comparison of Solar Panels

14 Power: Conclusions Li-Ion Cells: Worst case mass = 1298 kg Nominal mass (4 Days) = kg Nominal Volume (4 Days)= 0.25 m^3 Solar Panels: Mass = 4.16 kg Area= m^2

15 Thermal In AU units, difference between Earth and Moon is significantly small o Moon AU Earth AU = 1 AU Solar Flux I s = 1394 W/m 2 (at 1 AU) Stefan-Boltzmann Constant σ = 5.67 * 10-8 W/m 2 K 4 Total Power required for lunar crew module P int = 764 W Desired cabin temperature T = 293 K (Room Temperature)

16 Thermal Calculations Used full Stefan - Boltzmann equation and solved for T (Cabin Temperature) o o o A s = surface area exposed to sun A rad = total surface area T env = environment temperature

17 Thermal: Trans Lunar For the trans lunar case, we are assuming that the solar flux hits everywhere on the spacecraft expect for the base. o We make this assumption because this is the max surface area that the solar flux can hit on the spacecraft in free space assuming any orientation. We also assumed that the environment temperature in free space is approximately 0 C. o A reasonable assumption to make because (Cabin temperature) 4 >> (Free space temperature) 4

18 Thermal: Eclipse General For the eclipse conditions, we are assuming that the spacecraft is completely eclipsed during both the earth and moon orbits. During the earth orbit, we assume that the environment temperature is 280 K based on the solar flux at the earth's distance from the Sun. During the moon orbit, we assume that the environment temperature is 0 K because it is considered to be free space.

19 Thermal: Lunar Surface The Surface of the Moon has different temperatures at different location and time. Below is the most extreme temperatures to design for worst cases. Lunar Surface Temperature (K) Dawn 120 Polar 230 Dusk Sun Angle 370 Noon Equatorial 390

20 Thermal: Lunar Surface Lunar crew module will have different exposure to Sun at different times o Excluding the bottom surface area (no sun exposure) At Polar/Dusk/Dawn o 1/3 of surface area exposed At 45 Sun Angle o 1/2 of surface area exposed At Noon Equatorial o o All of surface area exposed Sun is directly above lunar crew module

21 Thermal: Coating Material Properties For initial thermal calculations, calculated all cabin temperatures with different coating o Different lunar surface temperatures o Different emissivities (ε) and absorptivities (α) Coating Absorptivity (α) Emissivity (ε) White Black Aluminum Polished Metals

22 Thermal: Lunar Surface Trade Study

23 Thermal: Emittance From our trade study of different coating materials, we found that aluminum coating (ε = 0.3 and α = 0.3) has the least variation of cabin temperature from the desired cabin temperature of 293 Kelvin. After doing thermal calculations for all the different conditions (trans lunar, eclipses, and moon surfaces at different times) with aluminum coating, we decided to have the lowest temperature on the trip be equal to the required cabin temperature of 293 Kelvin, so as to only utilize radiators and not any heaters. For us to meet this condition we found that we would need to have an effective emittance of 0.056

24 Insulation Trade Study

25 Thermal: Multi Layered Insulation We were able to achieve this condition by adding 5 layers of an 850-3M Mylar-Aluminum Backing insulation with aluminum coating on both sides to get an effective emissivity of so as to get the cabin temperature during the coldest situation, during the moon eclipse, to be 293 Kelvin. Similar to the effective emissivity, the effective absorptivity of aluminum decreases with more layers. Final absorptivity α = 0.05 and emissivity ε =

26 Thermal: Cabin Temperatures Cabin temperatures in different situations with multi layered insulation (MLI) Situation Temperature (Kelvin) Temperature ( Celsius) Lunar Surface: At Dawn Lunar Surface: At Dusk Lunar Surface: At Polar Lunar Surface: At 45 latitude Lunar Surface: At equatorial Trans Lunar (Free Space) Eclipse in Earth Orbit Eclipse in Lunar Orbit

27 Thermal: Radiator Will use Traveling Wave Tube Amplifier (TWTA) radiators with Optical Solar Reflectors (OSR) covering them o Total Mass = 31 kg o Emissivity of OSR: ε =.77 o Absorptivity of OSR: α =.06 Radiators will be positioned between top and middle thrusters

28 Thermal: Radiator Used the Stefan-Boltzmann equation to solve for the area required for OSR depending on the power generated by the spacecraft. Area of radiators 2 m 2 o o o Radiators will deploy away from the lunar surface at all times so that the environment temperature is reduced. Designed for the worst case condition that the OSR panels are facing the sun at all times Calculated for the OSR panels area to perform at room temperature for the cabin.

29 Thermal: Position of Radiators Side View of Radiators Top View of Radiators (when deployed)

30 Propulsion: Requirements Limited 6 DOF 1. Translational delta V = 50 m/s 2. Attitude Hold in Dead Band for Return 3. Overcome 500 Nm of Aerodynamic moments due to reentry (Pitch and Yaw) 4. Rotate Spacecraft 180 degrees in 30 seconds (Roll)

31 Propulsion: Attitude Control System The attitude control system will consist of 25 attitude thruster nozzles. All the nozzles will be recessed into the craft walls so that they will be protected from forces on the nozzle walls from drag forces and heating on re-entry. The nozzles near the heat shield need to be placed slightly higher up from the bottom because of the extreme heating of the heat shield region.

32 Attitude Control System Diagram 4 thrusters radially at the top for pitch/yaw adjustments in conjunction with radial thrusters on bottom 1 axial thruster at the top for translational movement along the z-axis

33 Attitude Control System Diagram 4 thrusters in the x-y plane around the center of gravity. For translational motion only - no moment generated because of placement.

34 Attitude Control System Diagram 16 thrusters (4 separate groupings) spaced equally radially. Each grouping has one radially, one axially, and two (opposite) azimuthally. Radial thrusters used for pitch/yaw motion and balance in the moment they produce, with thrusters on the top. Azimuthal thrusters used for roll movement. Axial thrusters for translational movement axially with the one on the top (calibrated to balance the power of the thruster axially on top.)

35 Propulsion: Translational Assumptions 1. delta V = 50 a. The total delta V required for all translational adjustments 2. Rocket Equation is a sufficient model 3. Mass and volume are primary considerations 4. Power and storage requirements are secondary considerations

36 Propulsion: Translational Selection I Cold Gas H2 He Methane N2 Air Argon Krpton Freon 14 SOA NTO-MMH 1 1=NASA-Document, See Sources Slide

37 Propulsion: Translational Selection II These Technologies were excluded: Type - Reason Electrical - High Power, Low Thrust Nuclear - High Mass Solid (Chemical) - Configuration Air Breathing (Chemical) - In Space Sails - Mission Design ED Tether - Mass and Volume

38 Propulsion: Translational Reasoning 1. Examine simplest case first (Cold Gas) 1. Examine State of the Art 1. Plot Mass versus Volume for a given system

39 Propulsion: Translational Mass versus Volume

40 Propulsion: Translational Conclusions I The best option is the SOA NTO-MMH This is due to: Isp=324 seconds Constituent densities at stored temperatures Oxidizer (NTO, Nitrogen Tetroxide) Fuel (MMH, Monomethyl Hydrazine) Earth Storable (liquid at ~290 K) Oxidizer to Fuel Ratio=1.65 Hypergolic (combusts on contact with each other) Utilized on the Space Shuttle RCS system

41 Propulsion: Translational Conclusions II For our translational propulsion requirements we will need: NTO: Mass = kg Volume = m^3 Tank (PMD) Mass = kg MMH: Mass = kg Volume = m^3 Tank (PMD) Mass = kg

42 Propulsion: Translational Conclusions III From NASA 2 the pressurization system with He to keep the fuel and oxidizer vapor stable is about 2 kg He per 55 kg (fuel+oxidizer) Thus: Mass He= ATM (Mass Opt. Tank) and 293 K Volume He= 1.09 x 10-4 m^3 Tank Mass He= kg

43 Propulsion: Dead Band Assumptions The amount of acceptable drift in roll, pitch, and yaw assumed for the dead band drift during the trip to Earth is 5 degrees in either direction from the desired path. It was also assumed that the lowest burn time for the thrusters is 0.1 seconds as limited by the solenoid valves that control the flow of propellant through the engine.

44 Propulsion: Dead Band Reasoning An angular velocity can be calculated from the torque that each set of thrusters delivers when fired, the moment of inertia about the axis of rotation, and the lowest burn time assumed.

45 Propulsion: Dead Band Reasoning Once an angular velocity can be determined as the result of an impulse bit delivered from a set of thrusters, the approximate amount of fuel consumed to correct the dead band drift over the entire trip to Earth can be calculated. After the first initial drift, the time before the next adjustment is completely dependent on the angular velocity generated by the impulse bit, which is extremely small.

46 Propulsion: Dead Band Data The amount of fuel required to correct for drift during the 3 day trip to Earth is negligible. The masses of fuel consumed were less than a gram of propellant total. Because the force requirements of the thrusters are so low as the result of strategic placement, the mass flow-rate of fuel through the engines is very small: 34.8 g/s for the roll thrusters and 44.1 g/s for the pitch and yaw thrusters.

47 Propulsion: Dead Band Conclusion The conclusion that can be drawn regarding fuel consumption for corrections to the dead band drift is that the craft barely drifts at all, and thus negligible amounts of propellant are consumed. This finding can be attributed to the use of a Isp propellant and a high moment of inertia on the craft, as well as a very strategic arrangement of thrusters around the spacecraft. Extraneous factors that were not included in this analysis justify applying a safety factor to the amount of propellant allocated for the translational propulsion to ensure enough fuel in case adjustments must be made to attitude.

48 Propulsion: Pitch and Yaw Assumptions In order to be able to counteract a moment of 500 N-m, two coordinated thrusters, positioned at the top of the craft and along the bottom, are fired. The two thrusters also generate equal forces, negating any translational velocity that might be imparted to the craft.

49 Propulsion: Pitch and Yaw Reasoning In order to calculate the required force to negate a 500 N-m moment during re-entry, the two thrusters must deliver combined moments that equal 500 N-m. The forces required to generate those moments are significantly lower due to the large moment arm for the thrusters positioned at the top of the vehicle.

50 Propulsion: Pitch and Yaw Data and Conclusion When taking this into consideration, each pitch/yaw thruster needs to be able to deliver 70 N of force. This also means that each thruster has a mass flow-rate of 44.1 g/s, as listed before. Such a low mass flow-rate is especially useful for a re-entry situation, where a long, continuous burn may be necessary to keep the vehicle stable.

51 Propulsion: Roll Assumptions The roll thrusters are positioned azimuthally in a ring as close to the base of the vehicle as possible without risking damage from reentry heating to give the highest possible moment arm. Each roll thruster is also assumed to deliver the same amount of force.

52 Propulsion: Roll Reasoning An equation from the lecture notes on propulsion was used to find the required torque to achieve the requirement of rotating 180 degrees in 30 seconds or less.

53 Propulsion: Roll Reasoning By applying a constant torque to accelerate the roll of the craft to 12 degrees/sec and an equal and opposite torque to bring the craft to rest, the vehicle will experience the smoothest possible acceleration from a constant burn.

54 Propulsion: Roll Data and Conclusion The total constant applied torque needed to execute a 180 degree roll maneuver in 30 seconds is N-m. When divided by the radius of the ring of roll thrusters, the required force for each thruster is very low: 27.7 N. This means that the mass flow-rate for the roll thrusters is 34.8 g/s of propellant. This is lower than the mass flow-rate for pitch and yaw control, and the 180 degree roll maneuver, when performed over 30 seconds, consumes half a kilogram of propellant.

55 Summation Mass Tabulation System Component Mass (kg) Mass (kg) Crew Module Power Li-Ion Cell 1298 Solar Panels 4.16 Thermal 30 Radiator 30 Propulsion Fuel (MMH) Fuel Tank Oxidizer (NTO) Oxidizer Tank RC Nozzles 64 He Mass (at 200 atm) 3.64 He Tank Total 2660

56 Sources Hutton, R. E. Lunar Surface Models. Tech. no. SP Washington, D.C: National Aeronautics and Space Administration, Print. Zhongmin, Deng. "Optimization of a Space Based Radiator." Applied Thermal Engineering31 (2011): Web Vasavada, Ashwin R. Near-Surface Temperatures on Mercury and the Moon and the Stability of Polar Ice Deposits. Publication no. Icarus 141. Los Angeles, CA: Academic, Print. "Thermal Control System Design." N.p., n.d. Web. 8 Nov <