ENAE483: Principles of Space System Design Power Propulsion Thermal System

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1 Power Propulsion Thermal System Team B4: Ben Abresch Jason Burr Kevin Lee Scott Wingate November 8th, 2012

2 Presentation Overview Mission Guidelines Project Specifications Initial Design Power Thermal Insulation Radiators Propulsion Moment of Inertia Propellant Mass Requirements Cold Gas System Summary

3 Mission Guidelines Power System, Reaction Control System, and Thermal Equilibrium Design for crewed spacecraft for lunar mission Crew of three for 10-day nominal mission (plus three contingency days) Use the crew and life-support systems designed for the previous design project Maximum gross mass 4795 kg

4 Project Specifications Power System provides electrical power to spacecraft for all mission phases to and from lunar surface Propulsion System capable of limited 6 DOF control o Translational ΔV of 50 m/sec o Attitude hold in dead band for three days (return to Earth) o Able to overcome entry aerodynamic moments of 500 Nm in pitch and yaw o Able to rotate spacecraft 180 in roll 30 sec on entry

5 Project Specifications (cont.) Thermal control system can maintain cabin temperatures in following cases: o Full sun (translunar) o Eclipse (Earth/Moon orbit) o Lunar surface dawn/dusk/polar o Lunar surface 45 sun angle (high latitudes/ midmorning or mid-afternoon) o Lunar surface noon equatorial

6 Choice of Initial Design (Shown without Power, Propulsion, Thermal Elements) We chose to use the design of Team A4 This design offered the most efficient use of space with extra volume specifically assigned for propulsion and thermal components

7 Power Requirements from Crew Systems System Power (W) E.D.C. CO 2 Scrubbing 300 Trace Contaminates Control 150 Water Distillation 156 Dehumidifier 500 O2 Storage 15.3 N2 Storage 4.2 H2 Storage 1.1 (From pg. 78 o f Team A4 Crew System Assessment) Crew System Power = 1.13 kw Assumed: Avionics Power = kw Total Power Requirement = 1.33 kw

8 Comparison of Power Choices Four possible power types were considered (Batteries, RPS, Solar, and Fuel Cells) Batteries cannot sustain enough power without having a mass on the order of thousands of kg, which is unacceptable A Radioisotope Power System will not be suitable because the power requirement is just out of the range of of an RPS system (<1kW) Solar cells would be unable to provide power during orbital eclipses and on the dark side of the moon, restricting possible missions Fuel cells are the best option for this design because of the ability to store a larger amount of energy for an extended amount of time while requiring a relatively small mass for the system

9 Fuel Cells Electrochemical system: 2H 2 +O 2 -->2H 2 O + e- System will generate potable water as well as store the energy required for the mission Using P [W]*T [hr]=energy [kwhr] found a required total energy of kwhr and converted to Joules Using the energy needed as the energy produced by the process the amount of water was found and used to find the total amount of reactants needed to sustain the required power for the duration of the mission

10 Mass Requirements Component Mass Requirements (Kg) Mass Rates (Kg/hr) Volume (cubic meters) Hydrogen Gas See tank Oxygen Gas See tank Reactor 22.4 n/a n/a Oxygen Tank 1.1 n/a Hydrogen Tank 1.6 n/a 0.18 Total mass Water Generated

11 Calculation Assumptions Created a Matlab script to calculate requirements based on a given power requirement Assumed that the reactor mass and volume could be scaled to the amount of power needed Assumed 1140 kg/m 3 and 71 kg/m 3 for LOX and LH2 densities to find volume of reactants required Also assumed perfect reaction to get products

12 Fuel Cell Products The generated water will be used in the thermal system as well as creating more potable water. Excess water will be dumped The exothermic reaction from the fuel cell reaction gives off the energy required to sustain a constant power of 1.33kW

13 Thermal Insulation The emissivity of the insulation surrounding the capsule was chosen in order to ensure an equilibrium temperature of at least 290 K while eclipsed in orbit The equilibrium temperature in eclipse was calculated by balancing the internal power with the power radiated away from the capsule and solving for the equilibrium temperature:

14 Thermal Insulation (cont.) This graph shows that an insulation with an emissivity of 0.03 will give the necessary equilibrium temperature in orbital eclipse

15 Thermal Insulation (cont.) Thermal insulation with an emissivity of 0.03 was achieved with MLI. The following equation determines the effective emissivity of several layers of mylar:

16 Thermal Insulation (cont.) This graph shows that 4 layers of insulation will give the necessary emissivity of 0.03

17 Radiators A total area of 20 square meters of flat panel radiators could be installed on the conical surface of the capsule By turning on and off 1 square meter portions of the radiator panels, the capsule s equilibrium temperature can be regulated in most situations based on the power balance:

18 Supplemental Heating In hot lunar conditions, the radiators are not enough. The max power the radiators can dissipate on the moon and still keep an equilibrium temperature of 298K is: This assumes that the sides of the cone radiate one third to the moon and the moon radiates as a perfect black body

19 Supplemental Heating (cont.) An ice sublimator system was added to dissipate additional energy when necessary. The power the sublimators must dissipate is: And therefore the mass rate at which ice must be sublimed is defined by:

20 Thermal Summary

21 Mass of Thermal Systems Part Mass Radiators (20 square meters) 220 kg Sublimator 20 kg Water for Sublimator (Worst case scenario: 7 days at lunar noon) Water Tank 208 kg 10 kg Total 438 kg

22 Moment of Inertia Assume the entire vehicle mass of 4795 kg is distributed in the hull for an approximate analysis Assume the thrusters will be located along the x-y centroidal plane Using the following equations for the moments of inertia about the tip and the parallel axis theorem, we find: I zz 3 10 m r2 I yy I xx 3 5 m 1 4 r2 h 2 I i,cg I ii mr i 2 I zz 8440kg m 2 I yy I xx 12600kg m 2 ENAE483: Principles 22 of Space System Design

23 Thruster Layout 16 thrusters (4 groups of 4) located around the craft on the x-y centroidal plane This de-couples the dynamics such that the action of a single thruster pair creates moments only about one axis Moment arm for all thrusters is 1.18 meters (from every axis they re not on) r arm = 1.18 m ENAE483: Principles 23 of Space System Design

24 Propellant Considerations Our propellant must meet the following criteria: 1. Low thrust to ensure long travel time between deadband impulses (and thus minimal fuel consumption) 2. High thrust to overcome reentry moments and rotations 3. Minimal mass requirements 4. Minimal volume requirements ENAE483: Principles 24 of Space System Design

25 Cold gas thrusters fit our model for what we need and are the focus of our research. The exit velocity, thrust, and mass flow rate (assuming the ambient pressure is zero) can be written as: e e e e thruster e e e thruster e e P V A T A V m P P M T V Cold Gas Propellant 25

26 Application of EOMS for Deadband Thrust occurs over a finite minimum interval t impulse o Limited by solenoid reaction speed, ~60ms Stable drift angular velocity is approximated by assuming the drift rates before and after the impulse are equal in magnitude: 2 I drift t impulse Time elapsed during a single drift is then: t drift deadband drift ENAE483: Principles 26 of Space System Design

27 Deadband Initial Velocity Determine the number of cycles required over the 3-day period of deadband control to be: N cycles t deadband t drift t impulse With t deadband = seconds Using the number of cycles, we can find the total mass required (per deadband axis): m deadband 2m thruster t impulse N cycles ENAE483: Principles 27 of Space System Design

28 Deadband Initial Velocity (cont d) Previous calculations assume small angular velocities. To keep this assumption, we need to consider slowing the craft from an arbitrary radial velocity By setting the final radial velocity to 0, we can find: I tslow initial And accordingly, the mass required is (per axis): m slow m thruster t slow ENAE483: Principles 28 of Space System Design

29 Improved Deadband Technique The deadband propellant mass is strongly linked to the finite time t impulse Using just two thrusters, we are required to burn for at least t impulse to remain balanced t impulse t impulse ENAE483: Principles 29 of Space System Design

30 Improved Deadband Technique (cont d) However, we can fire two more thrusters simultaneously, but at slightly longer times, say t impulse +t dt t impulse t impulse +t dt t impulse +t dt In doing this, we need to modify two of our equations t impulse ENAE483: Principles 30 of Space System Design

31 Improved Deadband Technique (cont d) Modified equations for the improved deadband technique: 2 I drift t dt t drift deadband drift N cycles t deadband t drift t dt m deadband 2m (2t t ) N thruster impulse dt cycles ENAE483: Principles 31 of Space System Design

32 Translational Velocity Requirements Using simple linear kinematics, we can find that: t trans m 2T vehicle V thruster m 2m vehicle thruster V V thruster And also: m m trans trans 2m m V vehicle thruster V thruster t trans ENAE483: Principles 32 of Space System Design

33 Re-entry Spin Requirement Assume constant thrust for 15 seconds each during which π/2 radians are traversed each (and then repeated in reverse). Applying our rotational EOMs: T T thrust thrust I r t arm 50.1N zz 2 spin ENAE483: Principles 33 of Space System Design

34 Re-entry Spin Mass Consumption We need to determine how much propellant is used to rotate the craft for all forces though. Applying our EOMs for rotational bodies, we find: t spin/ 2 m spin I T r thrust t 2m 2 spin/ 2 zz arm thruster ENAE483: Principles 34 of Space System Design

35 Re-entry Torque Requirement Applying a simple torque analysis (and assuming two thrusters contribute to working against the moment), we receive our first constraint on the thrusters: T thrust 500Nm 2 r arm T thrust 212.2N 500Nm m Thus, since this constraint is larger, our engines must produce more than N in thrust (based on their location) ENAE483: Principles 35 of Space System Design

36 Comparison of Cold Gas For consideration of our cold gasses, we ll use the below variable and consider them versus exit area Variables t impulse = 0.06 seconds θ deadband = π/18 rad P e = 2 psi P o = 300 psi t dt = seconds dθ initial /dt = π/36 rad/s T 0 = 300 K A e = varied Keep in mind, we re going to consider minimizing not just mass, but storage volume as well ENAE483: Principles 36 of Space System Design

37 Cold Gas Analysis Mass Budget ENAE483: Principles 37 of Space System Design

38 Cold Gas Analysis Volume Budget ENAE483: Principles 38 of Space System Design

39 Cold Gas Thruster Summary Freon-14 gas (stored at room temperature) represents the best performance while still being able to be stored Total Propellant Mass: 372.3kg o Total Propellant Mass with 20% Reserve: kg o Assumed Storage Mass: kg Total Storage Volume at 3000 psi (including Reserve): o m 3 o Approximate tank volume ~ m 3 (including gas) o Pressure drop via regulator, negligible mass ENAE483: Principles 39 of Space System Design

40 Design Overview Component Mass (Kg) Fuel Cell System 138 Thermal Control System 438 Attitude Control System 1340 Crew Systems 1303 Total 3219

41 Component Placement Thruster Quads Water Tanks Radiator Panels Propulsion Fuel Tanks Fuel Cell Fuel Tanks