PPT Design Project. ENAE483 November 8, 2010 Stef Bilyk, Kip Hart, John Pino, Tim Russell

Transcription

1 PPT Design Project ENAE483 November 8, 2010 Stef Bilyk, Kip Hart, John Pino, Tim Russell

2 Project Overview 1) Design the power system, reaction control system, and perform the thermal equilibrium calculations for the human spacecraft from the Crew Systems project. 2) Gross mass of the two projects = 4795 kg

3 Reaction Control System Specifications 1) Design a reaction control system for the human spacecraft chosen Must be capable of limited 6 DOF control Translational ΔV of 50 m/sec Attitude hold in dead band for three days (return to Earth) Able to overcome entry aerodynamic moments of 500 Nm in pitch and yaw Able to rotate spacecraft 180 in roll 30 sec on entry

4 Power System Specifications 1) Design a power system to provide electrical power to the spacecraft throughout the mission (with duration margin from last time) 2) Must support all mission phases LEO checkout Cis-lunar space LLO loiter Lunar descent and ascent Lunar surface operations Earth EDL

5 Thermal Control System Specifications 1) Design thermal control system (with radiator temperatures, sizes, and design locations on vehicle) to maintain cabin temperatures in the following cases: Full sun (translunar) Eclipse (Earth/Moon orbit) Lunar surface dawn/dusk/polar Lunar surface 45 sun angle (high latitudes/ mid-morning or mid-afternoon) Lunar surface noon equatorial - Can use supplemental radiators if necessary

6 Crew Systems Design Project Selection Selection: Team A10 Full list of detailed power requirements Sufficient volume for propellant, batteries, or other energy storage devices in design Total mass was well below maximum Most through descriptions of crew systems components Simplicity of atmosphere and other designs

7 Propulsion System Specifications Design an RCS for the given spacecraft 6 degree of freedom control 50 m/s of translational V 3 days of attitude deadband hold Reaction moment of 500 Nm during reentry Rotate spacecraft 180 under 30s for Earth entry

8 System Design RCS - 6DOF Control Quads of 4 thrusters spaced 90 apart Quads placed below floor of spacecraft cabin Quads are spaced 90 apart around the central axis of the spacecraft Firing a pair of thrusters produces pure rotation or translation coupled with rotation Purely rotational pairs can negate rotation associated with translation burns Orthogonality simplifies dynamics and controls

9 RCS - 50 m/s Translational V Force generated by 4 thrusters Vectors pointed towards tip of cone, normal to floor Consider orthogonal for initial design considerations Approx. with constant mass during V 16 kns impulse required from each thruster 60 mn minimum thrust

10 RCS - Attitude Deadband Deadband hold at ±5 Consider impulsive deadband maneuvers Spacecraft oscillates between deadband maxima Consider 14 impulses over 3 days 1.5 Ns impulse required from each thruster 0.12 mn minimum thrust

11 RCS - Reentry Moments Given moments of 500 Nm Moments experienced over 10 minutes of reentry Static equilibrium from opposing thrusters producing 156 N Assume a constant hold over entire reentry 94 kns impulse required from each thruster 156 N force applied by each thruster

12 RCS - Half turns under 30s Provide for 10 turns during mission Minimum angular velocity of 6 deg/s Assume constant thrust θ = π 900 rad s 2 t 2 θ = π rad 450 s 2 12 kns impulse required from each thruster 40 N force applied by each thruster

13 RCS - Thrust and Impulse Summary Requirement Thrust (N) Impulse (kns) Translational V Attitude Deadband Reentry Moments Half Turns 40 12

14 RCS - Analysis of Initial Design Large force requirements + Aerodynamic moments during reentry + Half turns Minimal force requirements + Translational V + Attitude Deadband Consider single system for moments and half turns

15 RCS - Large Force System Consider 160 N thrusters Thrusters controlled by solenoids, no throttle Assess force transmission to astronauts F trans = ( ) 160 N a trans = g 160 N cause acceptably low accelerations on crew

16 RCS - Large Force System Selection Single system providing 106 kns of impulse per thruster Each thruster produces 160 N of force Trade specific impulse with propellant mass m prop I sp = kg s log 10 m prop + log 10 I sp = 4 P jet = 1 2 TI spg o log 10 P jet = log 10 Tg o 2 + log 10 I sp Consider I_sp on the order of 100s of seconds Propellant mass on the order of 10s of kilograms Jet power around 1000 watts

17 RCS - Large Force Specifications Thruster provides 160 N Total impulse 106 kns + Delta-V 94 m/s + Thrust time 12 minutes Specific Impulse on order of 100s + Exit velocity on order of 1 km/s Jet power around 1 kw

18 Thrust Requirement

19 RCS - Large Force System Selection Narrow focus to chemical systems + Cold gas thruster + Monopropellant + Bipropellant + Hybrid system Consider first cold gas and monopropellant systems + Greater system reliability + Lower mass + Fewer moving parts Consider advanced systems if specifications are not met

20 RCS - Large Force from Cold Gas Specific impulse calculated from energy I sp = 1 g o 2η p ε Internal energy is in enthalpy + Consider ideal and calorically perfect propellant Replace energy with temperature and molecular mass terms Examine orders of magnitude log 10 T log 10 M = 6 Consider H 2, smallest propellant molecule log 10 T = 4 Total temperature required unacceptably high Cold gas thrusters are not appropriate for large force system

21 RCS - Large Force Monopropellant Decomposition of hydrazine peroxide Information available in references + Spacecraft Propulsion by Charles Brown Considering slope of spacecraft, select 40 lb thruster + can incline thruster 1 off the hull Minimum specific impulse of 115s Impulse time of 22 ms Thruster set mass of 22 kg + Thruster weight of 1.4 kg

22 RCS - Hydrazine Mass Requirement Reentry + Continuous thrusting, not pulsed + Endothermic reaction leads to inefficiencies + 220s of specific impulse + 44 kg of hydrazine required per thruster 707 kg of hydrazine required for reentry Half turns + Impulse bits of 22 ms and maximum off time of 62 ms + Specific impulse of 230s kg of hydrazine required per thruster 82 kg of hydrazine required for half turns

23 RCS - Large Force System Masses Thruster Mass 1.4 kg Propellant Mass 50 kg (0.049 m 3 ) Spherical Propellant Tank 23 cm diameter MER Tank Mass 15 kg Total system mass: 1060 kg

24 RCS - Minimal Force System Minimal force requirements + Translational V + Attitude Deadband Total impulse required is 16 kns Desire high thrust for V and low thrust for attitude + Need to throttle propellants to achieve both goals + Reconsider cold gas thruster Position thruster components above CTBs inside capsule Four groups of two thrusters positioned orthogonally

25 RCS - Delta-V Case Thruster V of 25 m/s + Thruster impulse of 32 kns Consider nitrogen gas system + Specific impulse of 70s + Propellant mass of 49 kg (negligible engine mass) Store nitrogen as a liquid (0.807 g/ml) + Requires m^3 of space + Spherical tank with radius of 24 cm Tank Mass kg tank from MER, include 10kg for throttling system Total Mass: 480 kg

26 Power Requirements Item Power Required (W) Intake and Supply Duct Fans 200 Cryogenic Vaporizer 6 4BMS 510 Water Distiller 73.5 Water Filter 1.5 EVA Suits 0 Avionics/Computer Systems 150 Total 941

27 Power Management and Distribution Traditional 28VDC system will be used PMAD system is only 85% efficient, this increases the total power generation requirements: 941watts Prequired, actual watts 0.85 Power needs to be provided for the entire duration of the mission at this minimum level

28 Power Generation and Energy Storage There are a number of different combinations of power generation and energy storage in space that may be appropriate for this mission: Solar arrays and batteries Fuel cells RTGs Batteries only The ideal system for a manned mission would be safe, relatively light, not overly complex, and well-tested.

29 Power Generation & Energy Storage: Fuel Cells Requires reactants, tanks, and reactor Maximum water mass savings of 21kg (the mass required in the previous design) kg/kw-hr reactants required for 7kW continuous fuel cells (Shuttle), assuming linear reduction in size with power requirement: m reactants W 0.339kg 24 hours ( )*( )*( )*( x 7000W kw - hr day days)*( kw) m reactor ( )*(255kg) kg ; m tanks. 128*( mreactants kg )

30 Power Generation & Energy Storage: Solar Arrays & Batteries Energy storage required for periods where the spacecraft is in shadow Mission will be designed to fit within a lunar day, so on the moon the panels will always be lit If the spacecraft is for some reason required to stay in LEO or LLO for at least one full orbital period, it will be in shadow. Power generation requirements increase in this case so that energy can be stored for operation during shadowed time period

31 Power Generation & Energy Storage: Solar Arrays & Batteries Worst case scenario: LEO Beta angle = 0 corresponds to: ~ 50 minutes (0.833 hr) sunlit ~ 40 minutes (0.667 hr) shadow Energy required during shadowed period: E stored 0.667hr *( W) W - hr Extra power generation required when lit: P extra W - hr hr W

32 Power Generation & Energy Storage: Solar Array Sizing Total power generation required from solar panels: P req, tot P extra P required w Using lightweight Si cells 17% efficiency 115 W/kg A arrays W m W (1394 )*(0.17) 2 m ; m arrays W W 115 kg 17.33kg

33 Power Generation & Energy Storage: Battery Sizing NiMH batteries used in conjunction with solar arrays 100 W-hr/kg, 80% Depth of Discharge m batteries W - hr (100 W - hr/kg) *(0.8) 9.23 kg NiMH batteries used alone (no recharging) m batteries only 24 hours W*( x days)*( ) day 100 W/kg

34 Initial Mass (kg) Power Generation and Energy Storage: Initial Mass Trade Study Batteries Only Fuel Cells Solar Arrays & Batteries RTG Days

35 Power Generation and Energy Storage Summary With increasing mission length, using only batteries quickly becomes extremely heavy and unreasonable. Fuel cells are relatively lightweight, but are expensive and complicated RTGs at this power level (based on those designed for Galileo) are heavy, generate a great deal of excess heat, and it is beneficial to avoid radioactive power generation for human missions

36 Power Generation and Energy Storage Summary Solar panels combined with batteries provide the lightest and simplest option for power generation and storage The required area of 8.4 m 2 is also very reasonable The mission will be timed so that the entire surface mission will be in sunlight The mass of batteries needed (10kg) to survive up to 40 minute periods of shadowing with only 50 minute periods of sun is well within the space and mass constraints of the capsule These batteries are also sufficient to provide power during Lunar landing, Earth re-entry, and any other large maneuvers that require either the stowage or redirecting of the solar arrays

37 Solar Array Design The arrays will have the ability to rotate about three axes The three rotational degrees of freedom will allow the arrays to track the sun over the entire course of the day so the power provided during the surface mission will be constant A nearly negligible amount of extra power is required to rotate the arrays very slowly There will be a moderate increase in structural mass to accommodate the rotating design The thin arrays will fold down to the sides of body during large maneuvers to reduce loads on the array structure The arrays are located near the top of the fuselage so that they will never be shadowed

38 Solar Array Design Summary Array Mass: kg + 10 kg (rotational structure) = 27.23kg Battery Mass: 9.23 kg Dimensions: 4.2 m x 1 m Total Array Area: 8.4 m 2 Array Thickness: 1.5 cm Sun tracking face has α p = 0.90 and ε p = 0.85 Opposite face covered in 20 layers of Mylar (ε 0.005) Can assume the solar panels are adiabatic on this side

39 Equatorial Noon Thermal Analysis

40 Internal Power Generated Assume that all of the electrical power used onboard the spacecraft is eventually converted into heat Also assume that humans produce heat P Humans P Electricity W ( 116 W )*(3 people) 348 W person P Internal W During high incident light periods, all internally generated power will need to be radiated During eclipse periods, not all internal power will be removed from the atmosphere in order to keep the cabin temperature at appropriate levels

41 Hull Properties Frustum with rounded bottom Total surface area: m 2 Coated in magnesium oxide paint α h = 0.09 ε h = 0.92 Chosen because it absorbs a small portion of heat from sun and emits a large portion of absorbed heat This keeps skin temperatures from getting to high when the spacecraft is in sunlight

42 Equatorial Noon Thermal Analysis Moon surface temperature: 380 K Solar panels incident area: 8.4 m 2 Hull incident area: 7.35 m 2 Q in = I panels hull sα p A incident + I s α h A incident + P int = 21,700 W Solar panels radiating surface area: 8.4 m 2 Hull surface area (not including bottom): m 2 Hull bottom surface area: m 2 Q out = A panels rad ε p σ T 4 eq T4 env + A top rad ε h σ T 4 eq + A bottom rad ε h σ T 4 eq T4 moon For Q in = Q out, T eq = 325 K T env

43 Radiators 325 K is too hot to maintain a comfortable cabin temperature Can add radiators to increase surface area Coated in magnesium oxide paint (for same reasons as hull) Can be oriented so they are oriented edge-on to the sun Eliminates any heat flow in from the Sun In order to get the hull temperature to 298 K need a total radiator surface area of 6.5 m 2 (3.25 m x 1 m) Thickness of 1.5 cm Fold up like an accordion which changes the effective radiating area and as a result the hull temperature

44 Polar/Dusk/Dawn Thermal Analysis

45 Polar/Dusk/Dawn Thermal Analysis Moon surface temperature: 180 K Solar panels incident area: 8.4 m 2 Hull incident area: 7.58 m 2 Solar panels radiating surface area: 8.4 m 2 Hull surface area (not including bottom): m 2 Hull bottom surface area: m 2 T eq = 282 K Assumes that all we choose to radiate all internal power

46 45 Sun Angle Thermal Analysis

47 Projected Area of a Cone The area of incidence of solar radiation on the spacecraft at a 45 sun angle is approximately the projected area of a cone. The total projected area is the sum of the elliptical region and the portion of the cone above it If the apparent height of the cone is less than the apparent height of the elliptical region, only the elliptical region needs to be considered. Images from Pennell, S., and J. Deignan. "Computing the Projected Area of a Cone." SIAM Review 31.2 (1989): 299

48 Projected Area of a Cone The area of the ellipse is given by: a e r cos( ) The area above the ellipse is given by: 2 a h (2hsin( ) u Where So h:height of cone * u ( u - * A tot * ( ) 3 r h ( hsin( )) - cos( ) ( r 2 r ) a h h 2 a r e :View angle 2 cot 2 ( ) 2 sin 1 above horizontal u ( r * ) u * r :Radius r 2 of ( u * ) cone 2 ) Equations from Pennell, S., and J. Deignan. "Computing the Projected Area of a Cone." SIAM Review 31.2 (1989): 299

49 45 Sun Angle Thermal Analysis Moon surface temperature: 215 K Solar panels incident area: 8.4 m 2 Hull incident area: 8.72 m 2 Solar panels radiating surface area: 8.4 m 2 Hull surface area (not including bottom): m 2 Hull bottom surface area: m 2 T eq = 286 K Assumes that all we choose to radiate all internal power

50 Full Sun Thermal Analysis Assuming Sun angle of 90 (hottest case) Solar panels incident area: 8.4 m 2 Hull incident area: 8.72 m 2 Solar panels radiating surface area: 8.4 m 2 Hull surface area (not including bottom): m 2 T eq = 280 K Assumes that all we choose to radiate all internal power

51 Full Shadow Thermal Analysis Solar panels incident area: 8.4 m 2 Hull incident area: 8.72 m 2 Solar panels radiating surface area: 8.4 m 2 Hull surface area (not including bottom): m 2 T eq = 162 K Assumes that all we choose to radiate all internal power

52 Mass Summary Component Mass (kg) Crew Systems Solar Arrays Batteries 9.23 Radiators Large Force Propulsion System 1060 Small Force Propulsion System 480 Total 2976

53 References Absorptivity & Emissivity table 1 plus others. Solar Mirror. Web. 08 Nov "High Accuracy Calculation for Life or Science." High Accuracy Calculation for Life or Science. Web. 08 Nov Pennell, S., and J. Deignan. "Computing the Projected Area of a Cone." SIAM Review 31.2 (1989): 299. Print. Soto, Laura T., and Leopold Summerer. POWER TO SURVIVE THE LUNAR NIGHT: AN SPS APPLICATION? Proc. of 59th International Astronautical Congress,. Vol. IAC-08-C