ppt design project - university of...

PPT Design Project

ENAE483 November 8, 2010 Stef Bilyk, Kip Hart,

John Pino, Tim Russell

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

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

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

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

• 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

Crew Systems Design Project Selection

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

RCS - 6DOF Control

System Design – 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

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

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

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


156 N force applied by each thruster

RCS - Half turns under 30s

– Provide for 10 turns during mission

– Minimum angular velocity of 6 deg/s

– Assume constant thrust

𝜃 =𝜋

450

𝑟𝑎𝑑

𝑠2


40 N force applied by each thruster

𝜃 =𝜋

900

𝑟𝑎𝑑

𝑠2𝑡2

RCS - Thrust and Impulse Summary

Requirement Thrust (N) Impulse (kNs)

Translational ∆V ≥0.032 16

Attitude Deadband ≥0.00012 0.0015

Reentry Moments 160 94

Half Turns ≥40 12

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

RCS - Large Force System

– Consider 160 N thrusters

– Thrusters controlled by solenoids, no throttle

– Assess force transmission to astronauts

𝑎𝑡𝑟𝑎𝑛𝑠 = 6 × 10−3 𝑔

160 N cause acceptably low accelerations on crew

𝐹𝑡𝑟𝑎𝑛𝑠 = (4 × 10−3) × 160 𝑁

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 𝑚𝑝𝑟𝑜𝑝𝐼𝑠𝑝 = 1.1 × 104 𝑘𝑔 ∙ 𝑠

log10𝑚𝑝𝑟𝑜𝑝 + log10 𝐼𝑠𝑝 = 4

𝑃𝑗𝑒𝑡 =1

2𝑇𝐼𝑠𝑝𝑔𝑜

log10 𝑃𝑗𝑒𝑡 = log10𝑇𝑔𝑜2

+ log10 𝐼𝑠𝑝

Consider I_sp on the order of 100s of seconds

Propellant mass on the order of 10s of kilograms

Jet power around 1000 watts

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

Thrust Requirement

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

RCS - Large Force from Cold Gas

– Specific impulse calculated from energy

𝐼𝑠𝑝 =1

𝑔𝑜2𝜂𝑝𝜖

– Internal energy is in enthalpy

+ Consider ideal and calorically perfect propellant

– Replace energy with temperature and molecular mass terms

– Examine orders of magnitude log10 𝑇 − log10𝑀 = 6

– Consider H2, smallest propellant molecule log10 𝑇 = 4

– Total temperature required unacceptably high

– Cold gas thrusters are not appropriate for large force system

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

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

+ 5.4 kg of hydrazine required per thruster

82 kg of hydrazine required for half turns

RCS - Large Force System Masses

– Thruster Mass

1.4 kg

– Propellant Mass

50 kg (0.049 m3)

– Spherical Propellant Tank

23 cm diameter

– MER Tank Mass

15 kg

Total system mass: 1060 kg

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

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 0.060 m^3 of space

+ Spherical tank with radius of 24 cm

– Tank Mass

+ 0.73 kg tank from MER, include 10kg for throttling system

Total Mass: 480 kg

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

• Traditional 28VDC system will be used

• PMAD system is only 85% efficient, this increases the

total power generation requirements:

• Power needs to be provided for the entire duration of

the mission at this minimum level

Power Management and Distribution

wattswatts

P 05.110785.0

941actual required,

• 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.

Power Generation and Energy Storage

Power Generation & Energy Storage:

Fuel Cells

• Requires reactants, tanks, and reactor

• Maximum water mass savings of 21kg (the mass required in the previous design)

• 0.339 kg/kW-hr reactants required for 7kW continuous fuel cells (Shuttle), assuming linear reduction in size with power requirement:

kW) (1.10705*days) (*)day

hours 24(*)

hr-kW

kg339.0(*)

7000W

W05.1107(reactants xm

; kg 40.32 kg)255(*)7000

05.1107(reactor m )kg(*128. reactantstanks mm


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

• Worst case scenario: LEO

– Beta angle = 0° corresponds to:

– Energy required during shadowed period:

– Extra power generation required when lit:



shadow hr) (0.667 minutes 40~

sunlit hr) (0.833 minutes 50~

hr-738.04W W)05.1107(*hr667.0stored E

W65.885hr 0.833

hr- W04.738extra P

• Total power generation required from solar panels:

• Using lightweight Si cells

– 17% efficiency

– 115 W/kg


Solar Array Sizing

w1992.71requiredextra totreq, PPP

; m 4.8

)17.0(*)m

W(1394

W71.1992 2

2

arrays A kg33.17

kg

W115

W71.1192arrays m

• NiMH batteries used in conjunction with solar arrays

– 100 W-hr/kg, 80% Depth of Discharge

• NiMH batteries used alone (no recharging)


Battery Sizing

kg 23.9(0.8)*hr/kg)- W(100

hr- W04.738batteries m

W/kg100

)day

hours 24(*days) (* W05.1107

only batteries

x

m

Power Generation and Energy Storage:

Initial Mass Trade Study

0

50

100

150

200

250

300

0 2 4 6 8 10 12 14

Init

ial M

ass

(kg)

Days

Batteries Only

Fuel Cells


RTG

• 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


Summary

• Solar panels combined with batteries provide the lightest and simplest option for power generation and storage

• The required area of 8.4 m2 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


Summary

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

• Array Mass: 17.23 kg + 10 kg (rotational structure) = 27.23kg

• Battery Mass: 9.23 kg

• Dimensions: 4.2 m x 1 m

• Total Array Area: 8.4 m2

• Array Thickness: 1.5 cm

• Sun tracking face has αp = 0.90

and εp = 0.85

• Opposite face covered in 20 layers

of Mylar (ε ≈ 0.005)

– Can assume the solar panels are adiabatic on this side

Solar Array Design Summary

Equatorial Noon Thermal Analysis

• Assume that all of the electrical power used onboard the spacecraft is eventually converted into heat

• Also assume that humans produce heat

• 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

Internal Power Generated

W05.1107yElectricitP

W348 people) 3(*)person

W 116( HumansP

W05.1455InternalP

Hull Properties

• Frustum with rounded bottom

– Total surface area: 33.765 m2

• Coated in magnesium oxide paint

– αh = 0.09

– εh = 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

Equatorial Noon Thermal Analysis

• Moon surface temperature: 380 K

• Solar panels incident area: 8.4 m2

• Hull incident area: 7.35 m2

𝑄 𝑖𝑛 = 𝐼𝑠𝛼𝑝𝐴𝑖𝑛𝑐𝑖𝑑𝑒𝑛𝑡𝑝𝑎𝑛𝑒𝑙𝑠

+ 𝐼𝑠𝛼ℎ𝐴𝑖𝑛𝑐𝑖𝑑𝑒𝑛𝑡ℎ𝑢𝑙𝑙 + 𝑃𝑖𝑛𝑡 = 21,700 𝑊

• Solar panels radiating surface area: 8.4 m2

• Hull surface area (not including bottom): 23.57 m2

• Hull bottom surface area: 10.195 m2

𝑄 𝑜𝑢𝑡 = 𝐴𝑟𝑎𝑑𝑝𝑎𝑛𝑒𝑙𝑠

𝜖𝑝𝜎 𝑇𝑒𝑞4 − 𝑇𝑒𝑛𝑣

4 + 𝐴𝑟𝑎𝑑𝑡𝑜𝑝

𝜖ℎ𝜎 𝑇𝑒𝑞4 − 𝑇𝑒𝑛𝑣

+ 𝐴𝑟𝑎𝑑𝑏𝑜𝑡𝑡𝑜𝑚𝜖ℎ𝜎 𝑇𝑒𝑞

4 − 𝑇𝑚𝑜𝑜𝑛4

• For 𝑄 𝑖𝑛 = 𝑄 𝑜𝑢𝑡, 𝑻𝒆𝒒 = 𝟑𝟐𝟓 𝑲

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 m2 (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

Polar/Dusk/Dawn Thermal Analysis

Polar/Dusk/Dawn Thermal Analysis







• 𝑻𝒆𝒒 = 𝟐𝟖𝟐 𝑲

• Assumes that all we choose to radiate all internal

power

45° Sun Angle Thermal Analysis

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

• The area of the ellipse is given by:

• The area above the ellipse is given by:

Where

So

Projected Area of a Cone

)(cos2 rae

))()(sin()(cos-))sin(()(

-)sin(2( 2*2**

12

2

3** uru

r

ur

r

huuhah

)(cot)( 222* rhh

ru

ehtot aaA

cone ofHeight :h horizontal above angle View : cone of Radius :r

Equations from Pennell, S., and J. Deignan. "Computing the Projected Area of a Cone." SIAM Review 31.2 (1989): 299

45° Sun Angle Thermal Analysis







• 𝑻𝒆𝒒 = 𝟐𝟖𝟔 𝑲


power

Full Sun Thermal Analysis

• Assuming Sun angle of 90° (hottest case)





• 𝑻𝒆𝒒 = 𝟐𝟖𝟎 𝑲


power

Full Shadow Thermal Analysis





• 𝑻𝒆𝒒 = 𝟏𝟔𝟐 𝑲


power

Mass Summary

Component Mass (kg)

Crew Systems 1311.5

Solar Arrays 27.33

Batteries 9.23

Radiators 87.75

Large Force Propulsion System 1060

Small Force Propulsion System 480

Total 2976

References

• “Absorptivity & Emissivity table 1 plus others.” Solar Mirror. Web. 08 Nov. 2012. http://solarmirror.com/.

• "High Accuracy Calculation for Life or Science." High Accuracy Calculation for Life or Science. Web. 08 Nov. 2012. http://keisan.casio.com/.

• 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.3.1.2.

http://solarmirror.com/

http://keisan.casio.com/



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