AIAA Design, Build, Fly Competition 2006/2007 Purdue Black

AAE 450 Spring 2010

AAE 4502/2/2010

Kathy Brumbaugh

612-860-2465

[email protected]

Chris Spreen

610-888-9521

[email protected]

AAE 450 Spring 2010

AAE 450 Spring 2010

Section 1 Presentations2/2/10

AAE 450 Spring 2010

Stephanie Sumcad2/2/2010

Trajectory Optimization

Airship Mission Design

1Stephanie Sumcad – 2/2

AAE 450 Spring 2010

Airship Descent Model

Trajectory Optimization 2Stephanie Sumcad – 2/2

V

W

θ

D

x

y

Airship Airship/Lander

Descent

Time (hr).94 5.02

Time (hr)

Altitu

de

(km

)

Airship Descent vs. Time Airship/Lander

Airship

gm

AvCr

refD

)21( 2

– Exponential density, gravity

gradient

– Post-aerocapture until 2km

altitude

Airship Cruising Cruising Altitude vs. Titan Winds

Path Options


Future Work Cruising path - Collecting data and dropping probes / lander

Descent - Position of entry about Titan, initial conditions, g-loads

requirements

AAE 450 Spring 2010

Altitude

(km)

vreq

(m/s)

vwind

(m/s)

vtotal

(m/s)

2 4.1654 4 8.16

1 4.1639 2 6.16

.1 4.1623 0.2 4.36

* Current airship capability is ~4m/s


2km

100m

Altitu

de

Altitu

de

Range

Range

OPTION 1

OPTION 2

AAE 450 Spring 2010

Backup Slides Mission Overview

– Lifetime: 6 months

– Circle Titan about 0° latitude 2x, 0° longitude 2x

– Deploy Lake Lander (90°N,0°W)

– Deploy 4 seismographs about Titan

({±90°N,0°W}; {0°N,0°W}; {0°N,180°W})


seismograph

drop point

cruising path

TITAN


AAE 450 Spring 2010

Backup Slides Vehicle Properties

Constants


m (kg) Aref (m2) CD

Airship 1694.6 78.5 .025

Airship and Lander 2139.1 78.5 .75


symbol Value units

Scale height H 40 km

Reference density (sea level) ρref 5.4 kg/m3

Reference gravitational accel. (sea level) gref 1.352 m/s2

Initial descent angle θ 0 deg

AAE 450 Spring 2010

Backup Slides Exponential Density

Model

H – scale height = 40km

ρref – reference density =

5.4kg/m3

rref – reference altitude = 0km


AAE 450 Spring 2010

Backup Slides Titan Wind Model

(Airship group members)

– 2m/s per 1km altitude


Altitude

(km)

vtotal

(m/s)

2 8.1654

1 6.1637

0.5 5.1629

0.4 4.9628

0.3 4.7626

0.2 4.5624

0.1 4.3623

Backup Slides Airship Descent Calculations – Matlab script

AAE 450 Spring 2010


Backup Slides EOM Integration Function

AAE 450 Spring 2010


Backup Slides - References

‗Aerodynamic Drag.‘ The Physics Hypertextbook.

http://physics.info/drag/

‗Terminal Velocity, gravity, and drag.‘ National Aeronautics and

Space Administration.

http://exploration.grc.nasa.gov/education/rocket/termvr.html

‗Post-Cassini Exploration of Titan.‘ Ralph D. Lorenz, 1999.

‗Engineering-Level Model Atmospheres for Titan and Mars.‘ C.G.

Justus, Aleta Duvall, and Vernon W. Keller.

AAE 450 Spring 2010




http://exploration.grc.nasa.gov/education/rocket/termvr.html

AAE 450 Spring 2010

Enoch Byers2/2/2010


Lake Lander Mission Design

Probe Comparison

Group Name (i.e.Trajectory Optimization)

Future Work Post landing a velocity of 2.2m/s

Many passes through the lake

-in the direction of wind then in the opposite direction

AAE 450 Spring 2010

Probe Mass (kg) Descent

Altitude(km)

Descent

Time

Landing

Velocity (m/s)

Lake Lander 500 2 5min 37.5s 3.46

Phoenix

Lander

350 124.885 >7m 2.1

Huygens 319 1270 2hr.

15min

<5

AAE 450 Spring 2010


Future Work

Refine Trajectory- Orbit insertion

Wind variation

Determine the best post landing trajectory

Backup Slide: Equations x = vi*t + .5*a*t2

F = ma

dv/dt = ((-.5*ρ*Cd*A*v2 )*sign(v)/M) - g;

A = Vf – Vi / t

AAE 450 Spring 2010

Backup Slides: Assumptions

Winds on Titan constant from the surface to an

altitude of 2 kilometers

For separate insertion wind is constant between

300-150km as well as 150-100km

Cd values assumed: .38(Parachute)

1.99(Probe Body)

1.55(3D heat shield drag)


AAE 450 Spring 2010

AAE 450 Spring 2010

Sid XIao2/2/2010


Escaping Earth

<1>

AAE 450 Spring 2010

Vehicle Centaur upper stage, PAM


Goal Escape 200km orbit with enough v∞

to achieve desired trajectory

Common Centaur SEC PAM

Total Mass(kg) 22744 2137

Dry Mass(kg) 1914 126

Isp(s) 450.5 286

Length(m) 11.7 1.83

Diameter(m) 3.05 1.24

See reference slide

<2>

http://upload.wikimedia.org/wikipedia/commons/e/e9/PAM-D_rocket_stage.jpg

Numbers A full burn of the Centaur with CBE payload

(From Andrew Rettenmaier) will give a v∞ of

7.8774km/s

To reach VEEJGA trajectory (v∞ = 4.7km/s, from

Chris Spreen) requires 68.18% of Centaur‘s fuel

Min. 4 PAMs to reach escape trajectory

For VEEJGA, need at least 6 PAMs


AAE 450 Spring 2010 <3>

Backup Slides


AAE 450 Spring 2010 <4>

Backup Slides


AAE 450 Spring 2010 <5>

Backup Slides


AAE 450 Spring 2010 <6>

Backup Slides

AAE 450 Spring 2010 <7>


00 lnsp

f

mv g I

m

2 2( )i c

p

v v vr

0

ft

t

Fv dt

m

Backup Slides

References

http://en.wikipedia.org/wiki/File:PAM-

D_rocket_stage.jpg

http://upload.wikimedia.org/wikipedia/commons/

b/be/Centaur_upper_stage_of_Atlas_V_rocket.jp

g

http://space.skyrocket.de


AAE 450 Spring 2010 <8>

http://en.wikipedia.org/wiki/File:PAM-D_rocket_stage.jpg



http://upload.wikimedia.org/wikipedia/commons/b/be/Centaur_upper_stage_of_Atlas_V_rocket.jpg



http://space.skyrocket.de/

References (con‘t)

Lockheed Martin. Atlas Launch System Mission

Planner‘s Guide

United Launch Alliance. Delta IV Payload

Planners Guide


AAE 450 Spring 2010 <9>

AAE 450 Spring 2010

Andrew Rettenmaier

Andrew RettenmaierFebruary 2nd, 2010

Structures – Orbiter Configuration

Mass and sizing of orbiter spacecraft core

structure

1

Preliminary Orbiter Configuration

Structure and Thermal Control GroupAAE 450 Spring 2010

Andrew Rettenmaier

2

Cylindrical envelope 4 m in

diameter and 3.5 m long

Core structure mass:

130 kg + 20% margin =

156 kg

Support structure mass:10% of hardware mass

~ 84 kg

Total of 241 kg

Ti-6Al-4V used in core

structure and rods for its

high strength to weight

ratio (wall thickness of 4mm)

Mass properties of preliminary

design (deployed wet)


Andrew Rettenmaier

3

Mass = 2500 kg

Center of mass* = [-51.5 -36.6 557.5] mm

Moments of Inertia** = [1477, 1551.2 1252.3] kgm2

Products of Inertia** = [-31.8 -58.9 39.5] kgm2

* Taken with reference to a cylinder 2m in diameter and 1.5m tall from center of end

** Taken with respect to CM

Does not take into account all systems currently+Z

+Y

+X

Back up Slides2

4


Andrew Rettenmaier

M = S/C point mass

L =Distance of S/C cg from interface

al = Lateral acceleration of LV

aa = Axial acceleration of LV

fa = Minimum axial s/c frequency

fl = Minimum lateral s/c frequency

M

Laa

al

Governing Equations:

[3][2]

[2]

A shorter orbiter structure is best – Titanium

material best option, though not necessary

5

Atlas VMass [kg]

Diameter [m]

Thickness [m]

Axial Mode [Hz]

Lateral Mode [Hz]

Msy Msu Msb

L=3.5m M=6000 - - 151 81 - - -

Al 6061 208 1.75 0.002 30.2 9.2 0.02 0.00 0.12

Ti 170 1.43 0.002 35.1 8.8 1.07 0.60 0.18

L=1.5m M=3450 - - 15 8 - - -

Al 6061 44 0.87 0.002 42.9 15.2 0.01 0.00 0.33

Ti 29 0.56 0.002 44.3 10.2 0.29 0.00 0.41

L=2m M=6000 - - 15 8 - - -

Al 6061 32 0.95 0.002 41.7 24.3 0.01 0.00 0.29

Ti 21 0.61 0.002 42.9 16.0 0.29 0.00 0.36

Axial Acceleration: 5.5g’s

Lateral Acceleration: 2g’s


Andrew Rettenmaier

Backup Slides

6


Andrew Rettenmaier

Launch Vehicle

Mass [kg]

Diameter [m]

Thickness [m]

Axial Mode [Hz]

Lateral Mode [Hz]

Axial Acceleration [g's]

Lateral Acceleration [g's]

Msy Msu Msb

Atlas V1 - - - 15 8 5.5 2 - - -

Al 6061 208 1.75 0.002 30.2 9.2 - - 0.02 0.00 0.12

Al 2090 187 1.64 0.002 30.7 8.8 - - 0.71 0.35 0.13

Ti 170 1.43 0.002 35.1 8.8 - - 1.07 0.60 0.18

Delta IV2 35 15 6 2

Al 6061 344 2.90 0.002 38.9 19.7 - - 1.53 1.50 0.14

Al 2090 299 2.63 0.002 38.9 17.9 - - 2.98 2.14 0.15

Ti 262 2.20 0.002 43.6 16.8 - - 3.46 2.45 0.21

Ariane V2 27 10 6 2

Al 6061 237 1.99 0.002 32.3 11.3 - - 0.33 0.32 0.13

Al 2090 217 1.90 0.002 33.1 11.0 - - 1.31 0.82 0.14

Ti 200 1.68 0.002 38.1 11.2 - - 1.86 1.21 0.20

Preliminary design von Mises stress

under allowable of 664 MPa

AAE 450 Spring 2010

Andrew Rettenmaier

7

Allowable yield stress of 664 MPa

Largest stresses in cross-beam supports of

653 MPa

5.5g

2.0g

3 translational

restraints applied to

base

Structure and Thermal Control Group

References

1.) “Atlas Launch System Mission Planner’s Guide”, Lockheed Martin

Corporation, Denver, CO, 2007.

2). J. J. Wijker, “ Spacecraft Structures”, Springer, Berlin, Germany, 2008

3). NASA SP-8007, Buckling of Thin-Walled Circular Cylinders, 1968.

7


Andrew Rettenmaier

AAE 450 Spring 2010

Chris Owen - Structures

Chris OwenFebruary 2, 2010

Structures – Lake Lander

Comparison of Structural Options

AAE 450 Spring 2010

Lander Structures

Chris Owen – Structures

Sub (Buckling due to pressure governs)

Radius (m) Length (m) Volume (m³) Hull Weight (kg)

0.2 1.6 0.2011 85

0.271 2.168 0.5002 175.3

0.342 2.736 1.005 310

0.4 3.2 1.6085 458

Floater (Volume, g-forces govern)

Radius (m) Length (m) Volume (m³)

Hull Weight (kg)

0.5 1.2 0.9434 107.4

0.8 0.47 0.9434 152

1 0.3 0.9434 179.3

2 0.075 0.9434 300.7

Floater w/ Probe (Hoop stress governs)

Radius (m) Length (m) Volume (m³)

Hull Weight (kg)

0.1 0.1 0.003 0.052

0.2 0.13 0.016 0.344

0.3 0.15 0.043 1.08

0.4 0.18 0.09 2.45

0.5 0.2 0.16 4.56

- Sub assumed to be a long tube

- Floater and probe assumed to be short,

cylindrical disks (hockey puck)

Floater-Probe Attachment

Chris Owen – Structures

Future Work Composite possibilities

Added support possibilities

Detailed stress/strain analysis

AAE 450 Spring 2010

Cable calculations

Carbon-Fiber AS2C (weakest available)

Probe Mass (kg)Probe Weights (Titan)

(N) Probe V (m³) Buoyant Force (N) Support Weight (N) Cross Sec. (cm^2)

75 101.25 0.16 84.8 16.45 0.0370

100 135 0.16 84.8 50.2 0.113

125 168.75 0.16 84.8 83.95 0.189

150 202.5 0.16 84.8 117.7 0.265

600 m Cable Weight (kg)

3.556

10.85

18.15

25.45

Backup Slides

Matlab codes


AAE 450 Spring 2010

Equations

AAE 450 Spring 2010


AAE 450 Spring 2010

Amanda Chastain

2 February 2010Interplanetary and Orbiter Propulsion

Nuclear Feasibility and Orbiter Propellants

AAE 450 Spring 2010

Nuclear Feasibility

Amanda Chastain - Propulsion

Specifications

Propellant Mass Fraction 0.65

Propellant Mass (kg) 4.39E+04

Vehicle Mass(kg) 8.72E+04

Payload Mass (kg) 5672

Initial Thrust Level (N) 2.57E+05

Mass Flow Rate (kg) 2.62E+01

Thrust Duration (s) 1.67E+03

Core Specifications

Power Core( W) 1.36E+03

Core Height (m) 2.09E+02

Core Radius (m) 1.15E+04

Core Volume (kg) 8.18E+10

Core Mass (kg) 1.31E+14 Not a viable choice!

Orbiter Propellants


Future Work Developing appropriate thruster system for

orbiter

AAE 450 Spring 2010

Monopropellant Isp(s) Propellant Mass (kg)

Hydrazine 230 361

Hydrogen Peroxide 190 442

HAN 262 315

Backup Slides- Nuclear Feasibility

Why choose Particle Bed Reactor?


AAE 450 Spring 2010

Assumptions:•Delta V = 4 km/s

•7 Element Reactor based Power required by core

Back up Slides - References Lawrence, T. J., Witter, K. J., & Humble, R. W. (1995). Nuclear

Rocket Propulsion Systems. In R. W. Humble, G. N. Henry, & W. J.

Larson, Space Propulsion Analysis and Design (pp. 443-507). New

York: McCraw Hill.

Osenar, M. J. (2004). A Comparison of Nuclear Thermal and

Nuclear Electric. Colorado Springs: U.S. Air Force Academy Dept of

Astronautics.

Peterson, C., & Kohlhase, C. (1997). The Cassini Mission to Saturn

and Titan. Retrieved January 31, 2010, from European Space

Agency: http://www.esa.int/esapub/bulletin/bullet92/b92kohlh.htm

AAE 450 Spring 2010

AAE 450 Spring 2010

David SmithDate 2/2/10

Interplanetary Propulsion

Solar Electric Propulsion

AAE 450 Spring 2010

Determining Mprop and ∆V

Propulsion Group

-Values are for 1 Thruster

- More propellant gives

higher ∆V, but increased

weight

- NEXT Thruster has a

450kg propellant minimum

Mprop (kg) Mtank (kg) Minert (kg) Msep (kg) Mi/Mf ∆V (m/s)

400 90.3604063 255.610746 745.971153 1.063943 2523.357

450 94.7933443 260.043684 804.837029 1.071884 2826.123

500 99.2013921 264.451732 863.653124 1.079815 3126.242

550 103.587825 268.838165 922.42599 1.087736 3423.758

600 107.955231 273.205571 981.160803 1.095645 3718.716

650 112.305701 277.556041 1039.86174 1.103543 4011.155

700 116.640952 281.891292 1098.53224 1.111431 4301.119

750 120.96242 286.21276 1157.17518 1.119309 4588.645

800 125.271318 290.521658 1215.79298 1.127175 4873.774

850 129.568684 294.819024 1274.38771 1.135032 5156.542

900 133.855414 299.105754 1332.96117 1.142877 5436.986

950 138.132291 303.382631 1391.51492 1.150713 5715.142

1000 142.4 307.65034 1450.05034 1.158538 5991.046

1050 146.659148 311.909488 1508.56864 1.166352 6264.73

1100 150.910274 316.160614 1567.07089 1.174156 6536.23

1150 155.153862 320.404202 1625.55806 1.18195 6805.576

1200 159.390346 324.640686 1684.03103 1.189734 7072.802

Thrust, Power and Volume

Propulsion Group

Future Work Work with trajectory to determine which configurations will get the spacecraft

there the fastest.

Determine the tradeoffs between speed, mass and power.

AAE 450 Spring 2010

Thruster PPU PMS,L PMS,H Gimbal Tanks

Power (W) 6390 0 15.9 4.3 0 0

Volume (m^3) 0.418 0.0312 0.00742 0.00317 0.095 0

Mass (kg) 12.7 34.5 3.1 1.9 6 Varies

Values per unit

# Thrusters Power (W) Thrust (mN) Vol (m^3)

1 6410.2 208 0.55479

2 12816.1 416 1.10958

3 19222 624 1.66437

4 25627.9 832 2.21916

Values for # Thrusters

Backup Slides Equations

Estimation for Xenon Prop

Estimation for Advanced Rigid/GaAs Solar arrays

Rocket Equation

Propulsion Group

AAE 450 Spring 2010

Backup Slides Given Values

Propulsion Group

AAE 450 Spring 2010

Isp (s) 4150

Lt (khr) 43.0

βs (kg/W) 0.0167

ρxenon (kg/m3) 5.984

mdot (mg/s) 5.12

Backup Slides Raw Data

Propulsion Group

AAE 450 Spring 2010

# Thrusters = 1


400 90.3604063 255.610746 745.971153 1.063943 2523.357

450 94.7933443 260.043684 804.837029 1.071884 2826.123

500 99.2013921 264.451732 863.653124 1.079815 3126.242

550 103.587825 268.838165 922.42599 1.087736 3423.758

600 107.955231 273.205571 981.160803 1.095645 3718.716

650 112.305701 277.556041 1039.86174 1.103543 4011.155

700 116.640952 281.891292 1098.53224 1.111431 4301.119

750 120.96242 286.21276 1157.17518 1.119309 4588.645

800 125.271318 290.521658 1215.79298 1.127175 4873.774

850 129.568684 294.819024 1274.38771 1.135032 5156.542

900 133.855414 299.105754 1332.96117 1.142877 5436.986

950 138.132291 303.382631 1391.51492 1.150713 5715.142

1000 142.4 307.65034 1450.05034 1.158538 5991.046

1050 146.659148 311.909488 1508.56864 1.166352 6264.73

1100 150.910274 316.160614 1567.07089 1.174156 6536.23

1150 155.153862 320.404202 1625.55806 1.18195 6805.576

1200 159.390346 324.640686 1684.03103 1.189734 7072.802

# Thusters = 2


400 90.3604063 418.889276 909.249683 1.062316 2461.071

450 94.7933443 423.322214 968.115559 1.070057 2756.662

500 99.2013921 427.730262 1026.93165 1.077788 3049.729

550 103.587825 432.116695 1085.70452 1.085508 3340.315

600 107.955231 436.484101 1144.43933 1.093219 3628.461

650 112.305701 440.834571 1203.14027 1.100919 3914.204

700 116.640952 445.169822 1261.81077 1.108608 4197.583

750 120.96242 449.49129 1320.45371 1.116288 4478.636

800 125.271318 453.800188 1379.07151 1.123958 4757.398

850 129.568684 458.097554 1437.66624 1.131618 5033.904

900 133.855414 462.384284 1496.2397 1.139267 5308.189

950 138.132291 466.661161 1554.79345 1.146907 5580.287

1000 142.4 470.92887 1613.32887 1.154537 5850.229

1050 146.659148 475.188018 1671.84717 1.162157 6118.049

1100 150.910274 479.439144 1730.34942 1.169768 6383.777

1150 155.153862 483.682732 1788.83659 1.177368 6647.444

1200 159.390346 487.919216 1847.30956 1.184959 6909.079

Backup Slides Raw Data

Propulsion Group

AAE 450 Spring 2010

# Thrusters = 3


400 90.3604063 582.167806 1072.52821 1.06077 2401.787

450 94.7933443 586.600744 1131.39409 1.068321 2690.534

500 99.2013921 591.008792 1190.21018 1.075861 2976.874

550 103.587825 595.395225 1248.98305 1.083392 3260.845

600 107.955231 599.762631 1307.71786 1.090912 3542.486

650 112.305701 604.113101 1366.4188 1.098424 3821.832

700 116.640952 608.448352 1425.0893 1.105925 4098.919

750 120.96242 612.76982 1483.73224 1.113417 4373.783

800 125.271318 617.078718 1542.35004 1.120899 4646.456

850 129.568684 621.376084 1600.94477 1.128372 4916.971

900 133.855414 625.662814 1659.51823 1.135835 5185.362

950 138.132291 629.939691 1718.07198 1.143289 5451.658

1000 142.4 634.2074 1776.6074 1.150734 5715.891

1050 146.659148 638.466548 1835.1257 1.158169 5978.091

1100 150.910274 642.717674 1893.62795 1.165595 6238.287

1150 155.153862 646.961262 1952.11512 1.173011 6496.508

1200 159.390346 651.197746 2010.58809 1.180419 6752.782

# Thrusters = 4


400 90.3604063 723.646336 1214.00674 1.059492 2352.681

450 94.7933443 728.079274 1272.87262 1.066884 2635.749

500 99.2013921 732.487322 1331.68871 1.074267 2916.504

550 103.587825 736.873755 1390.46158 1.08164 3194.983

600 107.955231 741.241161 1449.19639 1.089004 3471.22

650 112.305701 745.591631 1507.89733 1.096359 3745.25

700 116.640952 749.926882 1566.56783 1.103705 4017.107

750 120.96242 754.24835 1625.21077 1.111041 4286.824

800 125.271318 758.557248 1683.82857 1.118368 4554.432

850 129.568684 762.854614 1742.4233 1.125687 4819.962

900 133.855414 767.141344 1800.99676 1.132996 5083.445

950 138.132291 771.418221 1859.55051 1.140296 5344.911

1000 142.4 775.68593 1918.08593 1.147587 5604.389

1050 146.659148 779.945078 1976.60423 1.154869 5861.906

1100 150.910274 784.196204 2035.10648 1.162142 6117.491

1150 155.153862 788.439792 2093.59365 1.169406 6371.171

1200 159.390346 792.676276 2152.06662 1.176661 6622.973

Backup Slides References

Humble, Ronald W. Henry, Gary N. Larsen, Wiley J. Space

Propulsion Analysis and Design.

Patterson, Michael J. Benson, Scot W. NEXT Ion Propulsion System

Development Status and Performance. Glenn Research

Center, Cleveland, OH

Propulsion Group

AAE 450 Spring 2010

AAE 450 Spring 2010

Collin MorganFeb. 2, 2010

Launch Vehicle Propulsion

Upper Stage Propellant Requirement

Propellant Required to Leave LEO

Assumptions

– Centaur (CIII) Upper Stage

– Payload 6000 kg

– Circular parking orbit at 300 km

– Inert Mass Fraction of 0.095 (<1% error)

– 40% of upper stage propellant used to get into LEO

Plane

Change

[deg]

ΔV leave

[km/s]

Prop. Limit

[kg] Ref. [1]

Prop. for

LEO [kg]

(estimate)

Prop. to

leave LEO

[kg]

Fraction of

Prop. Used

0 3.20 20672 8268.8 7178 0.58

5 3.30 20672 8268.8 7535 0.60

10 3.58 20672 8268.8 8611 0.69

15 4.00 20672 8268.8 10444 0.84

20 4.52 20672 8268.8 13137 1.06

AAE 450 Spring 2010

Collin Morgan - Propulsion

Selection of Launch Vehicle Maximum allowable plane change while escaping LEO

(using 90% of total fuel):

– Centaur (RL10A engine)

• 14.7 deg

• 10,336 kg fuel required

– Delta IV Upper Stage (RL10B engine)

• 22.2 deg

• 13,600 kg fuel required

Future Work

– Incorporate 1st stage into maximum plane change

calculation

– Propellant required for aerocapture correction maneuver

AAE 450 Spring 2010


Earth

300 km parking orbit

Not to scale

Upper

Stage

Backup Slides ΔV Calculation

– Net inertial velocity to achieve circular orbit

– Escape velocity

– Required Change in velocity using law of cosines

vi Δv

vf

θ

where μ is Earth’s gravitational constant

and r is distance from center of Earth

Figure Based on Humble,

Henry, and Larson, Ref. 3Ref. [3]

AAE 450 Spring 2010


Propellant Required from Upper Stage to Reach LEO401 to GTO 551 to GTO 552 to LEO (450 km circ)

Event Time from liftoff [s] Time from liftoff [s] Time from liftoff [s]

MES1 251 267 263

MECO1 937 801 694

MES2 1502 1307 4230

MECO2 1730 1669 4242

Total burn time 914 896 443

Propellant Required to Leave LEO

Ref. [3]

AAE 450 Spring 2010


Propellant to Leave LEO Calc. Form of Ideal Rocket Equation

where Isp is specific impulse and finert is inert mass

fraction

AAE 450 Spring 2010


Calculations - Centaur

Centaur (CIII) RL10A

Plane Change [deg] delta V leave [km/s] Isp upper [vacuum] Isp feasibility [s] m_prop leave [kg] m_prop max Prop for LEOprop fraction

0 3.200114183 450.5 138.5837937 7178.2039 20672 8268.8 0.7472428

5 3.298961489 142.8644643 7534.861118 0.7644959

10 3.578478182 154.9691835 8611.11502 0.8165593

15 3.999147166 173.1866284 10443.61236 0.9052057

20 4.51905993 195.7019135 13136.60579 1.0354782

25 5.104701229 221.0636313 16903.29206 1.2176902

30 5.732233429 248.239472 22125.84481 1.4703291

60 9.728894849 421.3184529 308618.8513 15.329317

90 13.38140722 579.4937529 -110925.526 -4.9659794

AAE 450 Spring 2010


Calculation – Delta IV Upper

Delta IV Upper RL10B

Plane Change [deg] delta V leave [km/s] Isp upper [vacuum] Isp feasibility [s] m_prop leave [kg] m_prop max [kg] Prop for LEOprop fraction

0 3.200114183 462 138.5837937 6899.305598 27200 10880 0.6536509

5 3.298961489 142.8644643 7237.735428 0.6660932

10 3.578478182 154.9691835 8256.813261 0.7035593

15 3.999147166 173.1866284 9984.703381 0.7670846

20 4.51905993 195.7019135 12508.18018 0.8598595

25 5.104701229 221.0636313 16007.98796 0.9885289

30 5.732233429 248.239472 20806.34801 1.1649392

60 9.728894849 421.3184529 219164.8204 8.4575301

90 13.38140722 579.4937529 -120269.9361 -4.0216888

AAE 450 Spring 2010


References: [1] “Atlas Launch System Mission Planner’s Guide, Atlas V Addendum.”

International Launch Services., San Diego, CA. December, 1999.

[2] “Delta IV Payload Planner’s Guide.” United Launch Alliance,

Littleton, CO. September, 2007.

[3] Humble, J.W., Henry, G.N., and Larson, W.J., Space Propulsion

Analysis and Design. McGraw-Hill Companies Inc., New York. 1995.

Pages 12-76.

AAE 450 Spring 2010


AAE 450 Spring 2010

David Stone2/2/2010

Power Group, Orbiter Power Supply

Orbiter Power Choices

AAE 450 Spring 2010

Power Group

Comparison of Possibilities

Total Power Requirement: 942 W (Peak Power)GPHS-RTG ASRG

Initial Power Output [W] 290 155

Final Power Output (14 years) [W] 259 138

Operational Lifetime [years] > 14 > 14

Fuel Mass/Unit [kg] 7.561 0.8

Total Mass/Unit [kg] 55.5 23

Unit Dimensions [cm] 42.2 (D) x 114 (L) 72.5 (L) x 29.3 (W) x 41 (H)

Volume/Unit [m^3] 0.1594 0.08709

Power (EOM)/Unit Mass [W/kg] 4.6667 5.8261Operating Temperatures 1270 K - 566 K 923 K - 363 K

Number Req'd 4 7

Total Power output (BOM) [W] 1160 1085

Total Power output (EOM) [W] 1036 966

Total Mass [kg] 222 161

Total Volume [m^3] 0.6376 0.60963

Current Recommendation Use the option of 7 ASRG‘s

– Sufficient Power Output

– Significantly lower operating temps (923 vs. 1270 K)

– 61 kg less mass

– Much more efficient use of fuel

Future Work Power Scheduling

Feasibility of Scaling Power Designs

Cost Analysis for Power Design

AAE 450 Spring 2010

Power Group

Backup Slides

AAE 450 Spring 2010

Power Group

Power RequirementsSystem Requirement [W] Source

Science Payload 387 Payload Group

Communication 110 Comm. GroupAttitude Control 100 Attitude Group

Propulsion 10 Prop. GroupFlight Systems:

Structures and Mechanisms 15 TSSM Report

Thermal Control 33 TSSM Report C&DH 58 TSSM Report

Power Electronics Standby 20 TSSM Report

Losses (7%) 52 TSSM Report

TOTAL (20% Safety Factor): 942

Backup Slides Resources:

– “Basic Elements of Static RTG’s,” Rockwell International –Rocketdyne Division: T90d-29-121

– Callat, T., “Status of Skutterudite-Based Segmented Thermoelectric Technology Components Developed at JPL, 2006

– Shaltens, R.K., Wong, W.A., “Advanced Stirling Technology Development at NASA Glenn Research Center”

– “Space Radioisotope Power Systems: Advanced Stirling Radioisotope Generator,” October, 2008

– “TSSM: Titan Saturn System Mission - Final Report on the NASA Contribution to a Joint Mission with ESA.” 30 January, 2009

AAE 450 Spring 2010

Power Group

Backup Slides Power Decay Eqn:

P(t) = P0 – P0(1-0.5Δt/(half-life))

Δt = time duration (years)

Half-life = Half- Life of Pu-238 = 87.4 years

P0 = Initial Power Output

Most recent communications requirement

reduced from 840 – 110 W

Payload requirement raised from 247 – 387 W

AAE 450 Spring 2010

Power Group

AAE 450 Spring 2010

Alex BelshawFebruary 2, 2010

Power Group, Lake Lander Subgroup

Topic: Lake Heating Problem

AAE 450 Spring 2010

Insulation

Pyrogel (red)

Cryogel (blue)

Comes in sheets of 10 mm thickness

Hydrophobic

1 layer of each would add roughly 19 kg

Alex Belshaw: Power Group

Enclosed Insulation (Sub)

Insulation/Lake heating

With this insulation: 1330 W of heat into Lake

(Sub)

Causes ~ .0009 K/s Temp Change


Future Work Power management

New Lander configs – new mass

Refine Heat transfer analysis?

AAE 450 Spring 2010

Floater Configuration

Backup Slides

Assume rectangular prism for shape 2x1.5x0.5

m

Outer surface area of 9.5 m^2

ASRG dimensions: .725x.293x.410 m

Lay in a row along shortest dimension add 10%

of width between for structure

Surface area of Pyrogel around generators:

3.52 m^2


AAE 450 Spring 2010

Backup Slides

Pyrogel: 180 kg/m^3; Cryogel 130 kg/m^3

Mass = density*surface area*thickness

Q = Thermal Conductivity*deltaT*Surface Area/

Thickness (Q = K(Th-Tc)*A/L)

K = .05 W/(mK); L = .02 m; delta T = 560 K;

SA = 9.5 m^2


AAE 450 Spring 2010

Backup Slides

Thermo for Lake : U = Q

U = m*cv*(T2-T1)

Q = .05 W/(mK)*560 K* 9.5 m^2 / .02 m = 1330 W

T2-T1 = 1330 W / 800 kg / 1.8 kJ/(kgK) = .0009

K/s

Assume control volume of methane .25 m

around sub gives ~ 800 kg CV of methane


AAE 450 Spring 2010

Backup Slides

Other assumptions: thermal conductivity

constant, sub homogenous and other material

in it negligible for heat transfer,


AAE 450 Spring 2010

Backup Slides

References

http://www.aerogel.com/markets/industrial.html

http://www.engineeringtoolbox.com/methane-

d_980.html

Lake properties provided by Brandon Kan


AAE 450 Spring 2010

http://www.aerogel.com/markets/industrial.html

http://www.engineeringtoolbox.com/methane-d_980.html



AAE 450 Spring 2010

Power Group, Lake Lander Power Subsystem

Travis Ramp2 February 2010


Lake Lander Power Management

<1>

AAE 450 Spring 2010


Lake Lander Power Management

<2>

Maximum Power Requirement: 450-500W

System Power MassNominal

Lifetime

2 ASRG’s ~280 We ~45 kg 14 yrs

4 ASRG’s ~560 We ~90 kg 14 yrs

Same power requirements for all vehicle types

Batteries

Future Work Further Power Management Upon Vehicle Specification

Further Battery Analysis

<3>AAE 450 Spring 2010


Battery Type Amount Total Power Total MassLifetime

(Cycles)

Lifetime (After

Charge)

Quallion Lithium

Ion Rechargeable

Battery 11 259.2 We

1.82 kg

(4.01 lbs)> 100,000 ~ 3 hrs

Quallion Lithium

Ion Rechargeable

Battery 24 216 We

1.44 kg

(3.17 lbs)> 100,000 ~ 3 hrs

Quallion Lithium

Ion Rechargeable

Pouch Cell6 194.4 We

0.468 kg

(1.03 lbs)> 100,000 ~ 4 hrs

Backup Slides

System Power (W)

Instruments

Imager 7

Gas chromatographer/Spectrometer 75

Density/Pressure/Temperature Sensors 5.5

Sonar 4.8

Magnetometer 0.5

Microscope 10

Mud Scooper 15

Processing Device 20

Prop System 250

Communication System 50

TOTAL 437.8

Power Management Table and Summation



Backup SlidesBattery Comparison

System

Specific

Energy,

Wh/kg

Energy

Density,

Wh/L

Operating

Temp

Range, °C

Calendar

Life, yearsCycle life

Silver-Zinc 100 200 -10 to +25 <1 <100

Nickel-

Cadmium35 100 -10 to +25 >5 >30000

Nickel-

Hydrogen40 80 -10 to +30 5 to 10 >40000

Lithium-

Ion100 240 -30 to +40 4 1000

Ratnakumar Bugga*, Marshall Smart, Jay Whitacre and William West. Lithium Ion Batteries for Space Applications.



Battery Circuit Calculations P=VI (Power = Voltage*Current)

Series Circuit

– VT= V1+V2+V3…

– IT=I1=I2=I3…

Parallel Circuit

– VT= V1=V2=V3…

– IT=I1+I2+I3…

AAE 450 Spring 2010 Power Group, Lake

Lander Power Subsystem

<6>

Backup Slides



SOURCES• Dawson, Sandra. Battery Power.

http://inventors.about.com/library/inventors/blbattery1.htm

•Green, James. Planetary Systems Division Update.

http://nasascience.nasa.gov/researchers/sara/library-and-useful-

links/Green_PSS_June%20508%20final.pdf

• Quallion Batteries: http://www.quallion.com/sub-sp-main.asp

•Huygens Probe:

http://www.daviddarling.info/encyclopedia/H/Huygens.html

•Secondary Batteries: http://www.clyde-

space.com/resources/powerschool/power_storage/secondary_batteries

•Ratnakumar Bugga*, Marshall Smart, Jay Whitacre and William West.

Lithium Ion Batteries for Space Applications.

AAE 450 Spring 2010

Brent Kam-YoungFebruary 2, 2010

Attitude & Controls - Airship

Flight Control Hardware

AAE 450 Spring 2010

Brent Kam-Young (Attitude & Control)

List of Hardware & SchematicsHardware Quantity Consists of … Controls …

Stabilizer 4 Control Surface Pitch/Yaw

(Elevators/Rudders) Actuator System

Ballonet 2 Air pump Altitude/Pitch(Front/Back) Air scoop

Intake/Release valves

Designed by Alex Brunk & Brent Kam-Young

Important Values

Future Work Determine final mass/power/etc values

Determine values of and relation between Ballonet operation

and rise/descent time

Design Controllers for flight control hardware

Rudder & Elevators Quantity 4

materialaluminum 6061

density (kg/m3) 2700

mass (kg) 43.923

reference area (m2) 0.425

volume (m3) 0.10625

Actuators Quantity 4

mass (kg) 22

power (watt) 130Torque-peak (Nm) 240Torque-continuous (Nm) 77Cost $600-900

Resulting Moment on Stabilizer (Nm)

Deflection Angle of Stabilizer (degrees)

Airspeed (m/s) 10 20 30

1 0.1384866 0.272765 0.398756

2 0.5539464 1.091061 1.595025

3 1.2463793 2.454888 3.588806

4 2.2157855 4.364245 6.3801

5 3.4621648 6.819133 9.968906

6 4.9855173 9.819552 14.35523

AAE 450 Spring 2010


Backup Slides Force by Newton‘s 2nd Law:

• F = d/dt(m*v) = m*dv/dt + dm/dt*v

• dm/dt*v = rho*v^2*A*sin(deflection)

• F ≈ rho*v^2*A*sin(deflection)

– Assuming negligible change in velocity during operation of

rudders and elevators

– Assuming constant density of Nitrogen atmosphere

• Torque: τ = r X F = r*F*sin(θ)

Volume of Rudders & Elevators• Volume = 2*(reference area)*(thickness) +

(depth)*(thickness)*(perimeter length)

AAE 450 Spring 2010


Backup Slides References:

– " Actuation Systems From Teleflex

Aerospace," Teleflex, Teleflex Aerospace,

California, 2009.

[http://www.teleflexactuation.com/actuators.html. A

ccessed 1/31/2010.]

– " Motors and Servomotors," MOOG, Moog

Aircraft Group, New York, 2009.

[http://www.moog.com/products/motors-

servomotors/. Accessed 1/31/2010.]

AAE 450 Spring 2010


AAE 450 Spring 2010

Clara GarmanFeb. 2, 2010

Attitude Control, Group Contact

Control Hardware for Lake Lander Options

AAE 450 Spring 2010

Using a Submersible Rudders (3) and wings (2)

Sensors: gyroscopes, servo

motors, *sonar,

*imager with

lights

– Power : 20 W

– Mass : 12.4 kg (27.3 lbs)

– Volume : 4570 cm3 (279 in3)• Note these are for control mechanisms only

Attitude Control

Buoyancy

Wing

RudderWeight Gravity

Drag

Rudder

Thrust

AAE 450 Spring 2010

Using a Floater Rudder behind propeller or propeller gimbal

– Sensors: gyroscopes,

servo motors, *sonar,

*radar, *imager

• Power : 15 W

• Mass : 4.96 kg (10.9 lbs)

• Volume :1830 cm3 (112 in3)

Future work:• Determine center(s) of gravity for positioning of control

systems, create accurate FBDs to use

Attitude Control

a) rudderProp

b) gimbal

Thrust

Gravity

Drag

Wind

Weight

Buoyancy

Backup Slides

Floater data

– Sizing based on: using Al –ρ=0.098lb/in3,

approximate 4 in thickness, and equation from Dave

Gerr‘s ―The Nature of Boats‖:

• Rudder Area = 0.018*water line length*draft of hull=

0.018*2m*0.5m

– Assumed need of 3 servo motors (3W each) for

rudder (with 1 being redundant for malfunction)

– Assumed 3 gyros for pitch, yaw, roll

Attitude Control

AAE 450 Spring 2010

Backup Slides

Submersible data

– SEABASS configuration of surfaces assumed

applicable here

– Assumed same sized rudders and wings that are ½

the size of a floater rudder each

– Assumed 4 servos (1 redundancy in rudders and 1

in wings)

AAE 450 Spring 2010

Attitude Control

AAE 450 Spring 2010

Danny Glover2 / 2 / 2010

Aerodynamics & Heat Shielding

Entry Heating & Heat Shield Materials

AAE 450 Spring 2010

Heat flux model

Heat flux profile

Assuming

-linear velocity loss

-constant density

Aerodynamics & Heat Shield, Daniel Glover

Heat Shield Materials, Orbiter front shield

Text here


Future Work Refining heat flux model

Applying heat flux model to material selectionAAE 450 Spring 2010

Material Density Mass TRL Heat Flux

PICA 0.266 g/cm3 210 kg 9 ~11MW/m2

AQ-60 0.306 g/cm3 239.2 kg 9 ~2.5 MW/m2

SLA-561V 0.152 g/cm3 118.8 kg 6 3 MW/m2

Carbon-Carbon/Calcarb 0.214 g/cm3 167.3 kg 6 7 MW/m2

LI-900(Shuttle shielding

material)

1.44 g/cm3 1125.6 kg 9 Max heat

1800o K

Backup Slides Matlab code nD = 4; %nose diameter in meters

%velocity profile

v = 11600:-100:1818; %velocity in m/s

%denisty profile

d = 1.63e-4; %denisty in kg/m^3

pr = d/1.22522; %denisty ratio

%heat flux profile

qrad = 4.85e5.*(nD/6.09e-1).*(pr^1.65).*(v./3.048e3).^5.6; %heat flux for radiative heat

qcon = 374.6.*sqrt(6.09e-1/nD).*(pr .̂49).*(v./3.048e3).^3.81; %heat flus for convective heat

q = (qrad + qcon)/10000;

%heat flux plot

plot(v,q)

title('Heat Flux vrs Velocity')

xlabel('Velocity (m/s)')

ylabel('Heat flux (MW/m^2)')

Assumptions: linear velocity loss from 11.6km/s to 1.818 km/s, worst case density constant at 1.63e-4 kg/m3 from

lowest altitude in aerocapture.


AAE 450 Spring 2010

Backup Slides Sources Values for PICA heat flux from:

Stackpoole, M., Sepka, S., Cozmuta, I., and Kontinos, D., “Post-Flight Evaluation of Stardust

Sample Return Capsule Forebody Heatshield Material,” 46th AIAA Aerospace Sciences Meeting

and Exhibit,7 - 10 January 2008, Reno, Nevada.

Values for SLA-561V and Carbon-Carbon/Calcarb heat flux from:

Munk, M., and Moon, S., ―Aerocapture Technology Developments from NASA‘s In-Space

Propulsion Technology Program,‖ Planetary Science Subcomimittee Meeting, Oct. 3, 2008,

ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20070031974_2007032230.pdf.

Values for LI-900 from:

―Orbiter Thermal Protection System,‖ NASA Facts, National Aeronautics and Space

Administration, 2006. http://www-pao.ksc.nasa.gov/kscpao/nasafact/pdf/TPS-06rev.pdf.


AAE 450 Spring 2010

AAE 450 Spring 2010

Airship Fluid Dynamics

Michael BrodFebruary 2, 2010


Airship Buoyancy and Heating

AAE 450 Spring 2010

Latest Data

Use of Titan air preferable over

helium- not carrying gas leads

to less mass, smaller volume

Envelopes for temp range

between 200 and 800 K and N2:

mtotal = 1718 → 1728 kg

Vballoon = 375 → 625 m3

Mballoon fabric = 23 → 33 kg

Airship Thermodynamics

He

N2

Results

Mid-range waste heat from power source is 370

K, recommend using T=300 K, which

corresponds to mtotal=1723 kg, SAfabric=297 m2,

mfabric=28 kg

Current fabric has an area density of .094 kg/m2

Future Work Aerodynamic Shape

Thermal Analysis

AAE 450 Spring 2010

Airship Thermodynamics

Backup Slides- Nitrogen plots


AAE 450 Spring 2010

Backup-

Helium

AAE 450 Spring 2010


Misc Backupm_structures=114 kg

m_control=318.4 kg

m_comms=5.4 kg

m_instrument=69.3 kg

m_prop=90 kg

m_power=400 kg

m_probes=37.3*4 kg

m_thermal=548.3 kg

Altitude= 2 km

rho_atm=5.4*exp((0-z)/40)

AAE 450 Spring 2010


rho_atm=5.4*exp((0-z)/40) kg/m^3

T_atm=90 K

P_atm=rho_atm*Ru/MW*T_atm N/m^2

P_b=P_atm;

AAE 450 Spring 2010

Todsadol (Tep) Rungswang02/02/2010

Orbiter-Earth CommunicationLink Budget analysis and Cost to operate Deep Space

Network (DSN)

AAE 450 Spring 2010

Orbiter – Earth communication

Tep Rungswang - Communication

Todsadol Rungswang - Transmitting power as a function of transmitting dish diameter

Assumption: pointing error on spacecraft < 0.1

degree

Downlink: Bit rate 12 Kbps

Down link, transfer rate 12 kbps

Orbiter’s Antenna Pointing Error (deg)

Transmitting Power Required (Watt)

0.1 93

0.15 340

0.2 2150

For 4-m transmitting dish

Power: 110 W

Antenna’s size: 4m

Mass of the system: 110 kg

Figure by Charlie Tytler

Cost to operate DSN

Future Work Communication during Earth-Titan transfer

Redundancy of the system

AAE 450 Spring 2010

$0.00

$10.00

$20.00

$30.00

$40.00

$50.00

$60.00

$70.00

$80.00

$90.00

10 hrs 2.5 hrs 10 hrs 2.5 hrs

Co

st

of

Op

era

tio

n /

year

( M

illio

n

US

D)

70-m dish

34-m dish

2030 2050

Total cost for 2 years mission

34-m dish: $13.2 – $41.2 million USD

70-m dish: $52.6 – $165.0 million USD

Spreadsheet to calculate cost is uploaded on Google group site


Approximate number of hours / earth

day which orbiter can communicate

with earth = 10Jeremy Moon, Trajectory team

To add safety factor and reduce cost,

use 2.5 hours of communication / day

Backup Slides Mass Calculation

– Dish = 40 kg (Calculation based off SEABASS

mission)

– TWT Amplifier = 15 kg (SATCOM*)

– Transceiver = 15 kg

– Supporting structure / mount = 40 kg

– Total: 110 kg

AAE 450 Spring 2010


*www.satcomresources.com

Backup Slides Orbiter Downlink data calculation

AAE 450 Spring 2010


Science InstrumentsData sample Collected/yr

Pixels/Sample

Bits / sample Compress Factor Gbit/Yr

High Resolution Visible/Infrared Imager 1460 2.00E+06 12 5 7.008

High resolution thermal/Infrared Spectrometer 20000 2.00E+03 12 1 0.480

Radar Altimeter 365 8.64E+04 8 5 0.050

Surface Penetrating Radar 365 8.64E+04 8 5 0.050

Microwave Sounder 60000 1.00E+03 8 2 0.240

Ion/neutral Mass Spectrometer 20000 1.00E+03 12 1 0.240

radio occultation 3650 8.64E+04 8 5 0.505

magnetometer 1095 8.64E+04 8 5 0.151

ultraviolet spectrometer 20000 1.00E+03 12 1 0.240

accelerometer 3650 1.00E+03 8 5 0.006

Orbiter Status 36500 1.00E+03 1 1 0.037

Total Data from Orbiter 9.007Analysis by Tep Rungswang and Charles Tytler

Confirmed by Jacob Bills (payload group)

Backup slides

AAE 450 Spring 2010

Support Period Antenna Service Hours per No. Tracks No. Weeks Pre-, Post- Total Total Cost

No Name Size Year Track per Week Required Config. Time Reqd. for period

(

#) (description) (meters) (year) (hours) (# tracks) (# weeks) (hours) (hours) Real-Year

1 Scientific Data (34) 34BWG 2030 10 7.0 52.0 364.00 4004.0 15,502,502

2 Scientific Data (34) 34BWG 2050 10 7.0 52.0 364.00 4004.0 20,632,675

3 Scientific Data (70) 70 2030 10 7.0 52.0 364.00 4004.0 62,010,008

4 Scientific Data (70) 70 2050 10 7.0 52.0 364.00 4004.0 82,530,702

5 Scientific Data (34) 34BWG 2030 2.5 7.0 52.0 364.00 1274.0 4,932,614

6 Scientific Data (34) 34BWG 2050 2.5 7.0 52.0 364.00 1274.0 6,564,942

7 Scientific Data (70) 70 2030 2.5 7.0 52.0 364.00 1274.0 19,730,457

8 Scientific Data (70) 70 2050 2.5 7.0 52.0 364.00 1274.0 26,259,769

Cost calculation for operating Deep Space Network


Backup Slides Total downlink via orbiter

AAE 450 Spring 2010

Data Return to earth via orbiter (Gbit / Yr)

Orbiter 9.007 (Tep Rungswang)

Lake Lander 4.251 (Martin Czerep)

Airship 8.985 (Vic Strimbu)

Probe 11.004 (Charles Tytler)

Total 33.247

orbiter sees earth at 10 hours / 24 hours 41.67 %

Assume 1/4 of that available time to be conservative 10.416 %

Transmission period per year 3284789.76 sec

Required Transmission rate 10.123 Kbps

To add safety margin and possibly increase in science data return, I have done calculation for 12 Kbps downlink


Backup Slides Uplink to orbiter

– Much smaller data size since it only involves command/control

– Data rate = 1kbps (100 watts transmitting power for 6dB margin)

AAE 450 Spring 2010

ItemsData sample

send/yr Sample Bits Compress Factor Gbit/Yr

Command for orbiter 36500 1.00E+03 1 1 0.037

Command for airship 36500 1.00E+03 1 1 0.037

Command for Lake Lander 36500 1.00E+03 1 1 0.037

orbiter sees earth at 10 hours / 24 hours 41.67 %

Assume 1/4 of that available time to be conservative 10.416 %

Transmission period per year 3284789.76 sec

Required Transmission rate 0.034 Kbps

We have estimated 1 kbps

for uplink data rate to leave additional

room for extra data transfer.

This analysis does not pose much

constraint to our mission since the

transmitter on earth could increase the

transmitting power easily to

accommodate for bit rate required


Backup slides

Beamwidth = 21/(f*D) where f = frequency in GHz, and D = transmitting diameter in meter

Lpt = -12(e/Beamwidth)^2 ; Equation for calculating pointing error lost in dB where e = pointing offset in degree

Assumption made:

Deep Space Network (receiving antenna) is capable of pointing the antenna to within 1 – 2 milli degree [ http://deepspace.jpl.nasa.gov/technology/TMOT_News/aprl98/phr_drj.html ]

AAE 450 Spring 2010

Pointing Error Analysis

Figure by Charlie Tytler


http://deepspace.jpl.nasa.gov/technology/TMOT_News/aprl98/phr_drj.html

Backup Slides Line of sight with earth

AAE 450 Spring 2010

Approximate number of

hours / earth day which

orbiter can communicate

with earth = 10

Calculation done by Jeremy Moon

(Trajectory Group)


AAE 450 Spring 2010

Vic Strimbu2/2/2010

Communications

Airship Communications Analysis

AAE 450 Spring 2010

Vic Strimbu - Communications

Airship Antenna Location Options

On top of airship

– Uninhibited

Communications

– Melting of skin issue

In ―Gondola‖ under airship

– Increase transmission loss

to estimate communication

through skin

Location #1

Location #2

Updated Link-Budget Analysis

Future Work Risk Assessment

Redundant Systems for Emergencies

AAE 450 Spring 2010


Bit Error Rate Transmission Frequency Max Distance Antenna Efficiency Estimated Atmospheric Loss

1E-5 bps 2.04 GHz (S-band) 4820 km 0.6 8 dB

Location of Antenna In "Gondola" under Airship On Top of Airship

Type of Antenna and Dish Parabolic Dish Omni-Directional Parabolic Dish Omni-Directional

Data Transfer Rate (kbps) 5 5 5 5

Receiving Antenna Diameter on Orbiter (m) 0.13 0.13 0.13 0.13

Transmitting Antenna Diameter on Airship (m) 0.5 0.5 0.5 0.5

Power Consumption (W) 25 650 10 275

Estimated Mass in kg (including dish and hardware) 8 5 8 5

-Above Table is for Airship to Orbiter Transmission

Transmission from Orbiter to Airship

Data Transfer Rate (bps) Power Consumption (W)

17 5

-Table below is for Orbiter to Airship Transmission

Backup Slide – Bit Rate Calculation

AAE 450 Spring 2010


Science InstrumentsData sample Collected/yr

Pixels/Sample Bits / sample Compress Factor Gbit/Yr

High Resolution Visible/Infrared Imager 1460 2.00E+06 12 5 7.008

High resolution therman/Infrared Spectrometer 20000 2.00E+03 12 1 0.480

Radar Altimeter 365 8.64E+04 8 5 0.050

Ion/neutral Mass Spectrometer 20000 1.00E+03 12 1 0.240

radio occultation 3650 8.64E+04 8 5 0.505

magnetometer 1095 8.64E+04 8 5 0.151

MET Package 1825 8.64E+04 8 3 0.420

Nephelometer 8760 4.00E+03 8 3 0.093

Airship Status 36500 1.00E+03 1 1 0.037

Bit Rate Analysis Done By Vic Strimbu, Charlie Tytler, Tep Rungswang, and Matrin Czerep

Total Data from Airship

8.985

Orbiter Visibility: 7%

Transmission Period: 2207520sec Data obtained from instrument reports in pg 2-63, Figure 2-37

Total Transmission Data Rate: 4.07029173kbps Titan Explorer Flagship Mission Phase 1 Study Report - 2007Asking Transmission Data Rate: 5kbps

Backup Slide – Communication from

Orbiter to Airship

AAE 450 Spring 2010


• Transmission requires only Command and Control

• Therefore data rate is much lower (~17bps)

• Power requirements assumed to be very small

when compared to communication from airship to

orbiter (Estimated Max 5W)

Kyle Kennedy AAE 450 Spring 2010

Kyle Kennedy

Lake Lander Science Payload

Tuesday, Feb 2nd

Lake Lander Science Payload

Payload Requirements

*Mass 61.43 kg

*Power 218.7 W

*Communications 732.2 kbs

*Volume 0.0498

Traveling Range 1000km

*Includes a safety factor of 1.4


Mud Scooper Probe

Lake bed and area

surrounding lake may

have solids we are

interested in studying

Drill will allow to take

samples of harder

substances than the

mud

Future Work:

Placement of devicesKyle Kennedy AAE 450 Spring 2010

Backup — Breakdown of RequirementsDevice

Mass

(kg)

Power

(W)

Comm Req

(kbs)

Volume

(m3) width depth height

Imager with Lights 7 15 225 0.01

Gas chromatographer/spectrometer 23 75 4

Density/Pressure/Temperature Sensors 1.5 5.5 0.1 0.001

Sonar 2.53 4.8 4 0.001

Magnetometer 0.42 0.5 4 0.001

Microscope 1.5 10 200 0.02

Mud scooper 5 15 50 0.0001 4cm 4cm 8cm

Processing Device (est) 1 20 1 0.0001

TOTAL 41.95 145.8 488.1 0.0332

Optics information provided by Jeremey Voigt

Volume estimated for all parts

Comm Requriements provided by Martin Czerep


Backup — Kraken Mare Largest lake on Titan,

best opportunity to do

most science

Diameter: 1,170 km

Coordinates:

68oN,310oW

Images of Kraken Mare from Cassini


http://upload.wikimedia.org/wikipedia/commons/5/50/Kraken_mare.jpg

AAE 450 Spring 2010

Jacob Bills01/18/10

Science Payload

Orbiter Instrumentation

AAE 450 Spring 2010


Jacob Bills Science Payload

InstrumentMass (kg)

Power (W)

Volume (cm^3)

Pointing Restrictions

Placement Restrictions

Heritage Risk

Magnetometer 1.2 2.4 81.6 Yes Yes Yes Low

Accelerometer 0.24 0.15 28.8 No Yes Yes Low

Ultraviolet Spectrometer 6 4.8 760.8 Yes No (FOV) Alice Med

Ion/Neutral Mass Spectrometer 30 45.6 48360 Yes No (FOV) Cassini Med

Radar Altimeter 24 52.8 7586.4 Yes No (FOV) Yes Low

Microwave Sounder 120 150 144000 Yes No (FOV) Yes Low

Surface Penetrating Radar 18 72 54436.8 Yes No (FOV) Yes Low

HiRes Visible/IR Imager & Spectrometer*

48 36 420000

HiRes Thermal/IR Spectrometer*

42 24 360000

*Data taken from Jeremy Voigt


Power Budgeting Possible

– Max Power Range: 154-254 W

Jacob Bills Science Payload

Future Work

AAE 450 Spring 2010

Safety Factor: 1.2Kg W cm^3

Totals: 289.44 387.75 1035254.4

Detailed power budget

Finalize optional instrument layouts

Backup Slides

Science Payload

AAE 450 Spring 2010

INPUT: Base Values/No Safety Factor

InstrumentMass (kg)

Power (W)

Volume (cm^3)

Pointing Restrictions

Placement Restrictions Heritage Risk

Magnetometer 1 2 68 Yes Yes Yes Low

Accelerometer 0.2 0.125 24 No Yes Yes Low

Ultraviolet Spectrometer 5 4 634 Yes No (FOV) Alice Med

Ion/Neutral Mass Spectrometer 25 38 40300 Yes No (FOV) Cassini Med

Radar Altimeter 20 44 6322 Yes No (FOV) Yes Low

Microwave Sounder 100 125 120000 Yes No (FOV) Yes Low

Surface Penetrating Radar 15 60 45364 Yes No (FOV) Yes Low

HiRes Visible/IR Imager & Spectrometer*

40 30 350000

HiRes Thermal/IR Spectrometer*

35 20 300000

Backup Slides

Science Payload

AAE 450 Spring 2010

Placement Restrictions: Some instruments

located away from others, or specific

orientations

Magnetometer: Must avoid high-gain antenna

Accelerometer: Orientation must be known

FOV: Field of View

Risk: Low – Few, if any, changes in technology

Med – Slightly new technology may be required

Backup Slides

Science Payload

AAE 450 Spring 2010

Heritage: Has the equipment be used on various

missions

Pointing Restrictions: Instrument must be

accurately pointed (ex. In nadir direction)

Backup Slides

Science Payload

AAE 450 Spring 2010

References: APL, NASA, Titan Explorer Public Report, Jan. 2008

APL, ESA, JPL, NASA, TSSM Final Report Public Version, Jan. 2009

Yu, A., NASA, “INMS Engineering Technical Write-up”, California Institute of

Technology,

[http://saturn.jpl.nasa.gov/spacecraft/cassiniorbiterinstruments/instrumentscassiniinm

s/instcassiniinmsdetails/. Accessed 1/28/10]

Graham, S., NASA, “Aqua Project Science”, Goddard Space Flight Center,

[http://www.aqua.nasa.gov/about/instrument_amsu.php. Accessed 1/28/10]

AAE 450 Spring 2010

Adam CoulonFebruary 2nd

Science Payload

Airship Scientific Instruments & Surface Scoop

AAE 450 Spring 2010

Scientific Payload

Surface ScoopOption Mass (kg) Power (W) Dimensions(

cm)

Risk of

Failure

Abilities

1 0.4859 24 40 x 40 x 60 Low Scooping

2 1.599 62.4 40 x 40 x 60 Medium Scooping

Drilling

3 1.674 63.2 40 x 40 x 60 High Scooping

Drilling

Suction

Option 1 Option 2 Option 3

Updated Instrument Data

AAE 450 Spring 2010

Scientific Payload

Instrument Mass (kg) Power (W) Volume (cm^3)

Radar Altemeter 0.8 10.75 570.675

Mass

Spectrometer

12.2 20 690.201

Nephelometer 4 3 12376

Magnetometer 0.4 1.6 12.258

Meteorology

Package

1.4 1.6 18

Surface Scoop

(option 1)

0.4859 24 96000

Radio-Doppler 6 16 24580

Polarimter 2.5 5 675

HiRes V/IR

Imager

40 30 350000

Totals 67.7859 111.95 484922.134

Claw Backup Information

AAE 450 Spring 2010

Scientific Payload

Each arm of the claw was modeled as a 15cm

long aluminum cylinder with a diameter of 1cm.

There will be 3 arms total.

Claw_Mass=(15cm*π*(0.5cm)^2)*(2.7g/cm^3)*6t

otal cylinders=190.852 grams

Servo Motor Backup Information The servo motors used in the claw were modeled after

the DSMS-17 motor from Baldor

(http://www.baldor.com/support/literature_load.asp?LitNu

mber=FL1851)

Servo_Mass=0.295kg

Servo_Power=24 Watts

Dimensions: 5.59cm x 4.27cm x 5.83cm

AAE 450 Spring 2010

Scientific Payload

http://www.baldor.com/support/literature_load.asp?LitNumber=FL1851

http://www.baldor.com/support/literature_load.asp?LitNumber=FL1851

Drill Backup Information The drill was modeled as a right cone with a 5cm

diameter base and a height of 12cm.

Power and mass requirements were taken from

the LSD-1201PC drill from RYOBI

(http://www.ryobi.com.au/Assets/55/553afe3f-

93e4-41d5-aae2-6857d2b7ddc3.pdf)

Drill_Mass = 0.818 kg

Drill_Power = 14.4 Watts

AAE 450 Spring 2010

Scientific Payload

http://www.ryobi.com.au/Assets/55/553afe3f-93e4-41d5-aae2-6857d2b7ddc3.pdf









Suction Cups/Pump Backup Information A single vacuum pump will be used to create suction on the three

suction cups. This pump was modeled after the SP 135 FZ pump by

Schwarzer.

(http://www.schwarzer.com/pdf/SP_PUMPS_ProductBrochure.pdf)

Pump_Mass = 0.012 kg

Pump_Power = 0.8 Watts


3 suction cups modeled after the B2.5-42-NBR cup by Anver

(http://www.anver.com/document/vacuum%20components/vacuum

%20cups/cups-b2.5.htm)

Cup_Mass = 0.021 kg


AAE 450 Spring 2010

Scientific Payload

http://www.schwarzer.com/pdf/SP_PUMPS_ProductBrochure.pdf

http://www.anver.com/document/vacuum components/vacuum cups/cups-b2.5.htm




Backup Option Information Option 1: Claw + Servo Motor

Option 2: Claw + 2 Servo Motors + Drill

Option 3: Claw + 2 Servo Motors + Drill +

Vacuum Pump + 3 Suction Cups

AAE 450 Spring 2010

Scientific Payload

AAE 450 Spring 2010

AAE 450 Spring 2010

Project Managers2/2/2010

Kathy Brumbaugh

612-860-2465

[email protected]

Chris Spreen

610-888-9521

[email protected]

Agenda

AAE 450 Spring 2010

Time Task

12:25pm - 12:35pm Announcements & Brainstorming

12:35pm - 12:50pm

Technical Groups

• Mission Design Requirements review

• Brainstorm any issues/concerns

12:50pm - 1pm

Concerns

• Any data or information required from

other groups

• Pieces of the mission puzzle that are

missing

• General questions for the Team

AAE 450 Spring 2010

Project timeline

Project timeline

AAE 450 Spring 2010

Design Stages - DescriptionStage Description

Mission Design Requirements

(MDR)

•Identification of all subsystems for each group’s tasks.

•Tasks and steps that need to be accomplished for the subsystems identified

Preliminary Design Review

(PDR)

•Exploring all the options for each aspect of the design

•Generating numbers for each option

•Communicating these numbers between groups

**At end of PDR => Mission architecture possibilities have been identified.

Critical Design Review

(CDR)

•Project Managers and Group Leaders will meet to discuss all options and make decision

based upon group input and calculations.

•Choose one option for each aspect of the design (ex. Lake Lander = sub or boat)

•Investigate and generate numbers for options on that design aspect (ex. What type of

propulsion for the Lake Lander?)

•Work numbers in great detail

•Communicate between groups with detailed analysis and numbers for each option

**At end of CDR => Mission architecture is chosen (ex. Specific engine, launch vehicle and

numbers, fly-by or no fly-by, boat or submarine, lake lander & airship configuration)

Final Design Review

(FDR)

•Project Managers and Group Leaders meet to incorporate all group discussions into

choosing the best overall design including specifics for each category and aspect.

•Numbers and analysis are now completed in great detail.

AAE 450 Spring 2010

Project Managers

Announcements & Brainstorming Preliminary Design meeting:

– Purpose: to explore all options available to achieve

mission requirements

– When: Wednesday Feb. 10th, 8:30pm

ARMS 1028 (tentative)

– Who: Group Leaders

(NOTE: GLs will represent their technical group in

making sure that PM and APM understand all the

options and the pros/cons associated with each choice.

PM and APM will make the decision and present the

next Morning to the Team. Any concerns may then be

addressed)

AAE 450 Spring 2010

Announcements & Brainstorming Team name?

Team logo?

CAD tasks– Subgroup: Kyle Kennedy, Alex Brunk, Andrew Rettenmaier, Dan

Glover, Mike Iwanicki, Alex Belshaw, Adam Coulon, Brandon

Kan

– Others?

– SEP stage, subsystems (power, ASRGs), likely vehicles,

science equipment, airship+lake lander + orbiter transfer, + heat

shield, aeroshell, dynamic model for combined, separate entry

AAE 450 Spring 2010

Vehicle Subgroups Identify/document current

options for each technical

group

Determine/document initial

risk assessment associated

with each subsystem

Determine/document

numbers/information

needed from other vehicles

Identify/document any

current concerns

AAE 450 Spring 2010

Vehicle Subgroup example:

AAE 450 Spring 2010

Aerodynamics Attitude Communic-

ations

Payload Power Propulsion Structures Trajectory

Controller

option 1

Bit rate Instrument #1 Power option

#1

Prop idea #1 Material # 1

Frequency

List properties (max. stress, etc) and/or already-calculated

values (mass, power, volume, etc) associated with each

choice

What options have not been covered?

What are the risks associated with each choice?

What concerns do you have for the mission as a whole?

FRIDAY BY 5pm!!!

Mission Design Requirements

AAE 450 Spring 2010

Reference Requirement Type Modification Date

MDR – 001

Minimize the cost associated with delivering an orbiter, Airship

and Lake Lander vehicle to Titan within 10 years of launch

between 2020 and 2040.

Functional 1 1/28/10

MDR – 002

The orbiter vehicle must use aerocapture at Titan while the

Airship and Lake Lander vehicles use direct entry into the

atmosphere


MDR – 003The Lake Lander must be either a steerable boat or submarine

designed to probe the lake bottom and operate for 6 months. Functional 1 1/28/10

MDR – 004

The Airship must be steerable and propelled through the air by

either hot air, helium or hydrogen and be able to drop probes

into areas of interest to gain a three-dimensional analysis of

the internal structure of Titan in addition to gathering samples

from the surface.


MDR - 005

The orbiter vehicle must be placed into a nearly circular orbit

with the following specifications:

1. Inclination must be at least 85 degrees

2. The semi-major axis must be less than 3500 km

3. The orbiter must be operational for at least two years


MDR – 006

There must exist at least a 90% probability that:

1. The Orbiter will operate for at least two years

2. The Airship will operate for at least six months

3. The Lake Lander will operate for at least six months

4. The probes will operate for at least six months


Table 2, Mission Requirements

Agenda

AAE 450 Spring 2010

Time Task

12:25pm - 12:35pm Announcements & Brainstorming

12:35pm - 12:50pm

Technical Groups

• Mission Design Requirements review

• Brainstorm any issues/concerns

12:50pm - 1pm

Concerns

• Any data or information required from

other groups


missing


Concerns?• Any data or information required from

other groups


missing


AAE 450 Spring 2010

Documents

AIAA Design, Build, Fly Competition 2006/2007 Purdue Black