Power, Propulsion and Thermal Design Project ENAE483 Fall 2012

Transcription

1 Power, Propulsion and Thermal Design Project ENAE483 Fall 2012 Team B8: Josh Sloane Matt Rich Rajesh Yalamanchili Kiran Patel

2 Introduction This project is an extension of Team A2's Crew Systems Project The contents of this design report include: Reaction Control System Analysis Thermal Analysis Power Systems Analysis

3 Reaction Control System Analysis

4 Propulsion Systems Considered Propulsion systems that are fundamentally inadequate: Non-mass expulsion propulsion: sails and tethers are too costly by mass and volume Electric propulsion systems: require heavy power generation apparati and draw considerable power Nuclear thermal rockets: require massive reactors and radiation shielding Consider cold gas thrusters, which are an appealing propulsion alternative due to their simplicity and lack of need for heavy supplemental equipment

5 Cold Gas Thruster Propellants Propellants Density (kg/m3) Specific Impulse (s) Hydrogen Helium Methane Nitrogen Air Argon Krypton Freon Liquid Carbon Dioxide Determine which, if any, of these cold gas thruster propellants yields a feasible propulsion system in the context of the first design requirement

6 Requirement Breakdown: Translational ΔV Total translational ΔV of 50 m/s Higher specific impulse will give higher exit velocity Higher exit velocity will yield larger mass ratio Larger mass ratio means less propellant mass is required High Isp will yield less propellant mass required

7 Cold Gas Propellant Mass Trade Study Propellants Mass (kg) Hydrogen 81.9 Helium 135 Methane 210 Nitrogen 296 Air 319 Argon 410 Krypton 587 Freon Liquid Carbon Dioxide 351

8 Cold Gas Propellant Volume Trade Study Propellants Volume (m3) Hydrogen 243 Helium 123 Methane 23.7 Nitrogen 16.2 Air 14.5 Argon 9.92 Krypton 3.91 Freon Liquid Carbon Dioxide 0.360

9 Cold Gas Thruster Conclusion The propellant masses required to produce a translational ΔV of 50 m/s with a cold gas thruster are reasonable, but the smallest propellant volume is m3 for liquid carbon dioxide. Cold gas propellants occupy too much space within the volumetric constraints of this spacecraft. The ideal propellant has high Isp and high density to minimize both required mass and volume Seek to exploit the high density property of liquid propellant storage

10 Hypergolic Propellants Hypergolic propellants consist of a fuel and an oxidizer that ignite spontaneously when brought into contact Energy acquired from combustion allows for higher Isp Consider Monomethyl Hydrazine (MMH) and Dinitrogen Tetroxide (N2O4) as fuel and oxidizer, respectively Stored as liquids at room temperature: MMH density: 880 kg/m3 N2O4 density: 1450 kg/m3 Used by the Space Shuttle RCS and Orbital Maneuvering System and the Apollo Service Module RCS Use the R-4D Marquardt N2O4/MMH rocket engine as a reference for performance specifications Developed as attitude control thrusters for the Apollo Lunar and Service Modules

11 R-4D Rocket Engine

12 Mass and Volume of MMH and N2O4 What mass and volume of MMH and N2O4 are required to produce 50 m/s of translational ΔV? Mass (kg) Volume (m3) Monomethyl Hydrazine (MMH) Dinitrogen Tetroxide (N2O4) Total Significantly less costly by propellant mass and volume than any cold gas thruster propellant Determine whether hypergolic propellants MMH and N2O4 meet the remaining design requirements

13 Requirement Breakdown: Dead Band Attitude Control Attitude hold in dead band for three days Consider +/- 5 in pitch, yaw, and roll acceptable error Assume the following repeating process for pitch, yaw, and roll motion: Spacecraft begins at - 5 Impulsive burn causes spacecraft to drift in the positive angular direction Once at the + 5 error, an impulsive burn causes the spacecraft to drift in the negative angular direction Want to maximize the drift time (time spent between burns) in order to minimize total number of burns, total ΔV, and propellant mass required

14 Calculation of Drift Time Find the angular acceleration α in terms of thrust F, moment arm r, and moment of inertia I using torque equations (given that F is normal to r) Calculate angular velocity ω about the pitch, yaw, and roll axes by multiplying constant angular acceleration α by duration of burn tburn (assuming burn time is small compared to drift time) Impulse bit Ibit is defined as the product of thrust and burn time

15 Calculation of Drift Time (cont'd) Calculate the drift time tdrift between maximum error cases as where Θ=10.0 is the angular distance between the maximum error cases Goal is to maximize tdrift, which for a given Θ, I, and r, must be done by minimizing Ibit Minimum impulse bit for R-4D rocket engine is 15.6 Ns

16 Moment of Inertia Approximate the spacecraft as a cone with height h = 3.83 m, radius r = 1.79 m, and uniform mass m The center of mass is labeled as G in the figure Ixx = Iyy is the moment of inertia about the pitch and yaw axes, and Izz is the moment of inertia about the roll axis Moments of inertia will decrease slightly over time as propellant mass is depleted

17 Spacecraft with RCS The spacecraft will employ four equally-spaced clusters of four R-4D engines as its RCS This configuration allows for 6 degrees of motion 3 translational directions 3 rotational directions (yaw, pitch, and roll)

18 RCS Engine Cluster

19 Moment Arm In order to minimize coupling of translational and rotational motion, the center of each cluster of four thrusters is located in the plane of the spacecraft's center of mass, and two thrusters are fired in unison to achieve either translation or rotation The moment arm is the distance between the center of mass and the thrust vector. As shown in the drawing, r = 1.93 m

20 Moment Arm - Mass Moment arm dimensions: m2 cross sectional area 0.59 m length Number of moment arms 4 for the rockets 1 for the radiator Material Aluminum Density = 2700 kg/m3 Total mass = 5.4 kg

21 Propellant Mass Required for Attitude Hold Calculate the ΔV required for each attitude correction burn from the equation where Ibit is the minimum impulse bit and m is the mass of the spacecraft Calculate mass of propellant mp required for each burn from the following equations

22 Total Propellant Mass and Volume for Attitude Hold Iterate over total number of burns during three days, accounting for depletion of propellant mass, to calculate total mass of propellant required for yaw, pitch, and roll corrections MassYaw (kg) MassPitch (kg) MassRoll (kg) Total mass and volume of each propellant required to hold attitude in dead band Mass (kg) Volume (m3) Monomethyl Hydrazine (MMH) Dinitrogen Tetroxide (N2O4) Total

23 Requirement Breakdown: 500 Nm in Pitch and Yaw Overcome entry aerodynamic moments of 500 Nm in pitch and yaw With a moment arm of 1.93 m, the total thrust that must be generated to counter a moment of 500 Nm is F = 259 N

24 Dual Firing of Thrusters to Minimize Coupling In the frame of the image to the right, a counter-clockwise torque is developed by firing thrusters A and D in unison Because the thrusters are located in the plane of the spacecraft's center of mass, firing A and D in unison produces only rotation and no translation Each thruster has to provide half of the total required thrust, F = 130 N

25 Requirement Satisfaction and Propellant Mass The R-4D rocket engine produces 490 N of thrust, which by far surpasses the required 130 N, so this requirement is met Assume two thrusters are fired in unison for 30 seconds to correct each of pitch and yaw, resulting in a total thruster burn time of 120 seconds Mass and volume of propellant required to meet this requirement using the above assumption Mass (kg) Volume (m3) Monomethyl Hydrazine (MMH) Dinitrogen Tetroxide (N2O4) Total

26 Requirement Breakdown: Roll 180 in 30 Seconds Rotate spacecraft 180 in roll in 30 seconds or less For constant thrust, moment arm, and moment of inertia, angular acceleration as constant Calculate angular acceleration from the above expression, given that two thrusters are burning

27 Calculation of Burn Time Calculate an expression for angular velocity by integrating the angular acceleration over time Integrate angular velocity over time to get an expression for angular position. Set this expression equal to the desired rotation angle of 180 Solve for t, the burn time required to rotate the spacecraft 180 in a roll

28 Propellant Mass and Volume Required for Roll The spacecraft can roll 180 in 3.83 seconds, far less than the maximum 30 seconds The fast roll rate can be attributed to the high thrust of the RCS engines and the dual operation principle Total propellant mass and volume required for 180 roll Mass (kg) Volume (m3) Monomethyl Hydrazine (MMH) Dinitrogen Tetroxide (N2O4) Total

29 Total Propellant Mass and Volume Required Monomethyl Hydrazine (MMH) 92.1 Volume (m3) Dinitrogen Tetroxide (N2O4) Total Mass (kg) Include 50% more propellant to allow for additional attitude corrections or translational maneuvers during mission Monomethyl Hydrazine (MMH) 138 Volume (m3) Dinitrogen Tetroxide (N2O4) Total Mass (kg)

30 Total Propellant Mass and Volume Required (cont'd) Consider mass of tanks 0.5 kg / kg propellant Monomethyl Hydrazine (MMH) 138 Volume (m3) Dinitrogen Tetroxide (N2O4) Tanks 183 Total 549 Mass (kg) 0.315

31 Thermal Analysis

32 Surface Area of the Spacecraft Considering the conical section of the spacecraft only Half angle of 25 Total surface area = 33.7 m2

33 Cross Sectional Area Normal to the Sunlight For ϴ < 650, only the left half of the cone will be illuminated The power flux from the sun Qin is given by the surface integral of the left half of the cone 650 Is ϴ h 25 0 r

34 Cross Sectional Area Normal to the Sunlight The area of the left half of the cone normal to the incident ray can be determined by inspection for the following two cases ϴ = 00: Anormal = r*h ϴ = 900: Anormal = π*r2/2 To first order, Anormal can be approximated by the following equation without needing to evaluate the surface integral: for (ϴ < 650)

35 Preliminary Analysis for Various Cases Case Anormal (m2) Qin (W) Teq (K) Sun angle ( ) Full sun (translunar) Eclipse (Earth/Moon orbit) Lunar surface Lunar surface noon equatorial Assumptions Is = 1394 W/m2 (1 AU from the sun) No radiator Absorptivity = emissivity Tenv = 0 K Penv = 0 W Equations

36 Effect of Surface Coating at Eclipse Desired spacecraft temperature = 300K This is a plot of the equilibrium temperature of the spacecraft at eclipse At eclipse, there is no sunlight, so this is the coldest the spacecraft is going to be Also, no dependence on absorptivity, since no heat is being absorbed Conclusion: emissivity must be Any smaller, and even the coldest portion of the mission will be too hot without a radiator

37 Effect of Surface Coating at Noon on the Lunar Surface At noon on the lunar surface, the spacecraft will be as hot as it is going to be at any point in the mission Therefore, Teq for noon on the lunar surface should be at least 300K, otherwise a heater would always be needed Use a combination of absorptivity and emissivity along the Teq = 300K contour line This would minimize the radiator size

38 Iteration 1 - Choice of Surface Coating Surface coating optimization requirements Emissivity equal to or slightly larger than Choose emissivity and absorptivity such that Teq for noon is 300 K or slightly larger Iteration 1: Consider aluminum paint Absorptivity = 0.3 Emissivity = 0.28 Expectations based on preliminary analysis A radiator may still be needed with this configuration Noon would be slightly hotter than 300K, so a small radiator would be needed Eclipse would be about 100K below the desired temperature, so a heater would definitely be needed

39 Iteration 1- Detailed Analysis for Various Cases Case Sun angle ( ) Anormal (m2) Teq (K) Full sun (translunar) Eclipse (Earth/Moon orbit) Lunar surface Lunar 90 surface noon equatorial Assumptions Is = 1394 W/m2 (1 AU from the sun) No radiator Absorptivity = 0.3 Emissivity = 0.28 Tenv = 4 K Pint = 839 W Assume power consumption found during crew design enters the spacecraft as heat

40 Iteration 1 - Additional Power Needed to Maintain Heat Case Sun Teq (K) angle ( ) without added heat Qadded (W) Full sun (translunar) Eclipse (Earth/Moon orbit) Lunar surface Solve for how much additional heat Qadded needs to be added for Teq to equal 300K

41 Iteration 1 - Radiator Design Assume black body radiator Radiator will be a flat plate oriented such that: The plate is normal to the surface of the moon The sun's rays will hit the thin edge of the plate Therefore, the energy absorbed by the radiator is negligible What surface area is needed to cool the craft at noon? r Radiator surface area = 2.6 m2 Radiator radius = 0.64 m The radiator will be attached to the spacecraft with a thermal switch. Therefore, when the equilibrium temperature drops below 300 K, the radiator can be thermally isolated from the spacecraft, so heat will not be lost unnecessarily

42 Iteration 1 - Summary Iteration 1 summary: Aluminum paint Power needed: 3500 W Radiator surface area: 2.6 m2 Power needed is unreasonably high Iteration 2 To reduce power needed, use a surface coating with lower emissivity Select a corresponding value of absorptivity from the contour plot of equilibrium temperature at noon along the Teq= 300 K contour line

43 Iteration 2 - Choice of Surface Coating New material to consider: bare aluminum Emissivity: 0.03 to 0.10 Absorptivity: Specific values to use Emissivity: 0.07 Slightly above minimum value found from eclipse case Absorptivity: 0.03 Point on contour plot of Teq at noon slightly above 300 K This meets both optimal cases at the two extremes Need to determine whether bare aluminum satisfies intermediate cases as well

44 Bare Aluminium Surface Coating

45 Iteration 2 - Detailed Analysis for Various Cases Case Sun angle ( ) Teq (K) Full sun (translunar) Eclipse (Earth/Moon orbit) Lunar surface Lunar surface noon equatorial Assumptions No radiator Absorptivity = 0.03 Emissivity = 0.07 Tenv = 4 K Pint = 839 W

46 Iteration 2 - Heater Power to heat spacecraft Crew Systems: 839 W Additional power needed: 244 W for eclipse case If only 50% of the power used by the crew systems is converted to heat, then the heater needs to provide 664 W Vornado EH Personal Vortex Heater Power range: W 21.3 x 21.3 x 20.3 cm 0.54 kg Vortex heater not only generates heat but also circulates air throughout the cabin

47 Iteration 2 - Radiator Design The radiator is again designed to be able to handle the lunar noon surface case Radiator surface area = 0.49 m2 Radiator radius = 0.28 m Radiator mass = 17.3 kg Assuming material has the same density as aluminum Thickness =.005 m With this new surface, the radiator will need to be used intermittently to maintain 300 K temperature for the full sun case

48 Radiator

49 Vehicle With Radiator

50 Iteration 2 - Summary Iteration 1 summary: Aluminum paint Power needed: 3500 W Radiator surface area: 2.6 m2 Iteration 2 summary: Bare aluminum Power needed: 244 W Radiator surface area: 0.49 m2 Factor of 10 decrease in power needed Factor of 5 decrease in radiator surface area The design found on iteration 2 will be used for our final thermal design configuration

51 Power Systems Analysis

52 Power Systems Considered RTGs Radioisotope thermal Photovoltaic arrays Fuel cells Batteries Photovoltaic arrays with 24-hour battery backup Eliminated RTGs and Radioisotope thermal because of the potential danger that nuclear energy can pose to the astronauts

53 Power System Trade Study

54 Power System Trade Study Analysis Cannot use only solar arrays because the spacecraft will be in eclipse for part of mission With solar arrays eliminated, for a 13-day mission duration, fuel cells are the optimal power system by mass Even the most high density batteries are far too massive for use on the spacecraft Photovoltaic arrays with batteries to power the spacecraft during eclipse would still be more massive than fuel cells at 13 days

55 Power System Calculations Total of 526,032 W-hrs required for the full 13 day mission Power required by system: 839 W for crew system 184 W for rocket engine valves (at peak) 663 W for thermal systems during eclipse Assumed peak power requirement for propulsion and thermal systems during the whole mission

56 Power System Calculations (cont'd) 3 Component Mass (kg) Volume (m ) LH LH2 tank LOX LOX tank Reactor Total Assumptions: LH2 tank/lh2 ratio is LOX tank/lox ratio is Nominal reactor mass can be scaled by a factor of how much power is required

57 Propellant and Power Systems Storage Space

58 Power, Propulsion and Thermal Summary Propulsion Hypergolic propellants MMH and N2O4 Mass: 549 kg m3 Thermal Bare aluminum surface coating Vornado heater: 663 W Radiator surface area: 0.49 m2 Power systems LH2 and LOX kg m3

59 References Donabedian, Martin, and David G. Gilmore. Spacecraft Thermal Control Handbook. El Segundo, CA: Aerospace, Print. "Vornado EH Personal Vortex Heater." Amazon.com. N.p., n.d. Web. 07 Nov "Vornado Personal Heater." Bed Bath and Beyond. N.p., n.d. Web. 07 Nov "Civil and Structural Engineering Design & Construction Tutorial Resources Site." Center of Gravity & Mass Moment of Inertia of Homogeneous Solids Tutorials. N.p., n.d. Web. 07 Nov "Encyclopedia Astronautica N2O4/MMH." N2O4/MMH. N.p., n.d. Web. 07 Nov "Encyclopedia Astronautica R-4D." R-4D. N.p., n.d. Web. 07 Nov "John F. Kennedy Space Center -Â KSC Fact Sheets and Information Summaries." John F. Kennedy Space Center -Â KSC Fact Sheets and Information Summaries. N.p., n.d. Web. 07 Nov