spring 2004 aae450: slide 1 introduction brady kalb aae 450 – spring 2004 project homer humans...

Spring 2004 AAE450: Slide 1

Introduction

Brady Kalb

AAE 450 – Spring 2004

Project HOMERHumans Orbiting Mars for Exploration and Research


Homer Heavy Lift Launch Vehicle

44 m

89 m57 m

2nd Stage 3rd Stage

Chris Ulrich, Chris Krukowski, Frank Hankins, Nikolaus Ladisch, Marina Mazur, Matt Maier


HOMER LAUNCH VEHICLES

Initial Mass (kg)

Final Mass (kg) Propellant Mass (kg)

1st Stage Main Core

2,940,000 2,680,000 465,000

2nd Stage Main Core

2,480,000 461,000 2,020,000

3rd Stage 344,000 238,000 106,000

Strap-on Boosters 1,710,000 203,000 1,500,000

HEAVY LIFT LAUNCH VEHICLE MASS BREAKDOWN

Chris Ulrich, Chris Krukowski, Frank Hankins, Nikolaus Ladisch, Marina Mazur, Matt Maier


CRV: Aerodynamic Stability

a=144deg

Front (1)

Aft (0)

)()(0 CGMRCXCGMRCZmm zzCxxCCC

Equation used in Trim line calculations:

5 10 15 20 25 300

50

100

150Static Margin vs. Various Mach

Mach #

Sta

tic M

argi

n (%

)

Alpha=140Alpha=145Alpha=150

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1-1

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1Static Margin Mach 29.5

Xcg

(delX/length)

SM

(%

)

alpha = 110.1365alpha = 115.1365alpha = 120.1365alpha = 125.1365alpha = 130.1365alpha = 135.1365alpha = 140.1365alpha = 145.1365alpha = 150.1365alpha = 155.1365alpha = 160.1365alpha = 165.1365alpha = 170.1365alpha = 175.1365alpha = 180.1365

Rebecca Karnes


CRV: LES Sizing and Components

Boost Protective Cover

Launch Escape Tower

Rocket Structure• Launch Escape Motor• Pitch Control Motor• Tower Jettison Motor

Heather Dunn


CRV: LES and Parachute Mass

Component Property Value Drogues Number 2 Diameter (each) [m] 5.5 Area (each) [m2] 23.5 Main Parachutes Number 3 Diameter (each) [m] 33.6 Area (each) [m2] 888.0

Component Mass (kg)

Launch Escape Tower 517

Launch Escape Motor 2132

Boost Protective Cover 430

Pitch Control Motor 23

Tower Jettison Motor 50

Total 3152

Parachute Recovery System Launch Escape System

Heather Dunn


Transport Vehicle

Thrusting Mode after leaving Earth

Devin Fitting, Dave Goedtel, Ben Toleman, Debanik Barua


Transport Vehicle

Storage view with airlock

Devin Fitting, Dave Goedtel, Ben Toleman, Debanik Barua


Transport Vehicle

• Aerocapture Mode– Radiators retracted

– Comm. Antenna Retracted

– Vehicle collapsed


Human Factor Mass Summary

Component # of ItemsUnit Mass[kg/unit]

Total Mass[kg]

Comments

1st Floor Total 34 - 1280 See Table D‑2 for list

2nd Floor Total 20 - 1600 See Table D‑3 for list

3rd Floor Total 27 - 3420 See Table D‑4 for list

4th Floor Total 1 - 400 See Table D‑5 for list

Stored Items Total 22,500 - 16,900 Includes 11,600 kg of consumablesSee Table D‑6 for list

Other Items Total 4 - 12,700 Includes 11,800 of kg consumablesSee Table D‑7 for list

Installation Margin for Zero g

- 0.4 14,500

Total 50,800

Total with 5% Growth 53,300

(HF Consumable Mass) 23,400

(HF Dry Mass) 29,900

Pat Nelson, Steve Blaske, Theresa McGuigan, Wade McMillan


Major Components Contained on the First Floor


1st Floor Component # of ItemsUnit Mass[kg/unit]

Total Mass[kg]

Comments

Bed 4 46 184

Washing Machine 1 100 100

Dryer 1 60 60

Desk 4 15 60

Chair 8 5 40

Shower 1 75 75

Sink 1 8 8

WCS 1 112 112 Waste Collection System

Multi-gym 1 200 200

Stepper 1 136 136

Treadmill 1 150 150

Gym Equipment 1 25 25

Table 1 15 15

Couch 1 45 45

TV 4 10 40


Major Components Contained on the Second Floor


2nd Floor Component # of ItemsUnit Mass[kg/unit]

Total Mass[kg]

Comments

Microwave 2 35 70

Dishwasher 1 40 40

Sink 1 15 15

WCS 1 112 112 Waste Collection System

Small Sink 1 8 8

Med Suite 1 1000 1000

Bed 1 55 55

Desk 1 15 15

Table 1 15 15

Chairs 5 5 25

TV 4 10 40

Scientific Payload 1 200 200 Not Much! (i.e. biomass growth chamber, biogen water recycling)


Major Components Contained on the Third Floor


3rd Floor Component # of ItemsUnit Mass[kg/unit]

Total Mass[kg]

Comments

Console 5 130 650 Includes chair for console

Table 2 15 30

Chair 7 5 35

Mainframe 2 200 400

Large TV 1 30 30 Command TV

Work Table 1 20 20

TV 2 10 20

Airlock 1 400 400 Backup Unit

WPA 2 658 1320

OGA 2 140 280

4BMS 2 120 240


Stored Components



Total Mass[kg]

Comments

Food* 3600 2.3 8280 2nd Floor

Cleaning Supplies* 900 0.25 230 Evenly divided between floors

Cooking Supplies 4 5 20 2nd Floor

Bathroom Supplies* 3600 0.05 180 1st and 2nd

Backup Bathroom Bags 3600 0.25 900 1st and 2nd

Personal Hygiene Kit 4 1.8 8 1st Floor

Hygiene Supplies* 3600 0.075 270 1st Floor

Clothing 4 90 360 1st Floor

Recreation Items 1 1000 1000 1st Floor

Personal Items 4 50 200 1st Floor

Vacuum 1 13 13 1st Floor

Disposable Wipes* 3600 0.3 1080 2nd Floor

Trash Bags* 3600 0.05 180 Evenly divided between floors

Operational Supplies 4 20 80 Includes diskettes, ziplocks, tape…(Evenly Divided between floors)


Stored Components (Continued)


Component # of ItemsUnit Mass[kg/unit]Total

Mass[kg] Comments

Restraints 1 100 100 For zero g environment

Hand Tools 1 300 300 Primarily 3rd floor

Test Equipment 1 500 500 3rd Floor

Other Maintenance Equipment

1 1000 1000 3rd Floor

Photography 1 120 120 1st Floor

Fire Suppression 4 5 20 Evenly divided between floors

EVA Tools 1 123 120 3rd and 4th

Manuerving Unit 2 35 70 4th Floor

EVA Suits 4 135 540 Primarily 4th

Med Consumables* 1 500 500 2nd Floor

Crew 4 70 280 Evenly divided between floors

Water Tank Spares 1 329 329 Hab Exterior

Waste Spare 1 56 56 3rd Floor

Atmosphere Spare 1 130 130 3rd Floor


Stored Components


Major Components Contained on the Fourth Floor

4th Floor Component # of Items

Unit Mass[kg/unit]

Total Mass[kg]

Comments

Airlock 1 400 400 Primary Unit

Other Components


Total Mass[kg]

Comments

Water Tanks 1 204 204 Allotted Tank Mass

Water* 1 10199 10,199

Air Tanks 1 743 743 Allotted Tank Mass

Total Gas* 1 1566 1566


Air Subsystem

Atmosphere Composition and Pressure

Gas Pressure [kPa]

p(O2) 19.50

p(CO2) 0.12

p(N2) 50.38

Total Pressure 70.30

Air Subsystem Breakdown

Component Mass [kg]

Total Gas 1,600

Mechanical Systems 500

Tanks 700

Spares 300

Air Subsystem Design Values

TotalsValue Unit

Total Mass 4,500 kg

Total Volume 7.0 m3

Total Power 2.6 kW



Waste & Water Subsystem

WCS Specifications

Spec. Value Units

Mass 112 kg

Volume 0.55 m3

Power 375 Watts

Daily Water Budget

Spec. Value Units

Water Input 118 kg/d

Water Output 119 kg/d

Percent Recycled 90 %

Mass Recycled 107 kg/d

Difference between Required and Recycled

10.5 kg/d

Water Mass for MissionSpec. Value Units

Required Mass for Recovery of Recycling Losses

9600 kg

Mass of Initial System Charge 120 kg

5 % Margin for Leakage / Spills 490 kg

Total Mass of Water 10,200 kg

Packing Factor 1.02

Total Loaded Mass of Water 10,400 kg

Volume of Water [0.001 m3/kg] 10.4 m3

WPA SpecificationsSpec. Value Units

Mass 658 kg

Volume 2 m3

Power 915 Watts



Artificial Gravity

Gravity Gradient relative to 9.81 m/s2

Floor Transit Gravity Martian Orbit Gravity

1st 1 0.38

2nd 0.92 0.34

3rd 0.83 0.29



Volume Comparisons


Comparison

Volume per Person[m3/p] Comments

Small Office 15 Habitable volume per person

Recreational Vehicle 27.5 Habitable volume per person

Naval Submarine 145 Pressurized volume per person

Skylab 100 Pressurized volume per person

Mir 124 Pressurized volume per person

ISS 142 Pressurized volume per person


10.5 m

10.084 m

2.58 m

Habitat Module

Debanik Barua, Masaaki Atsuta, Jamie Krakover, Rachel Malashock, George Mseis, Daniel Nakaima


Storage Module

Doors for CRV/Landers

10.5 m

10.084 m



Effect of Thickness on Hoop Stress



Buckling Analysis



Column Configuration and FEM Analysis

1.00 m

0.02 m

R 0.20 m

R 0.20 m

0.10 m

1.50 m

2.00 m

2.58 m

0.75 m

R 0.30 m

Max. von Mises Stress = 9.65×107 N/m2

Max. Principal Stress = 9.74×107 N/m2

Max. Displacement = 1.36×10-4 m Mass = 916.34 kg



Brace Configuration and FEM Analysis



Max. Displacement = 6.25×10-4 m Mass = 65.80 kg



Floor Configuration and FEM Analysis



Max. Displacement = 1.40×10-4 m Mass = 9.76×103 kg



CRV and Lander Holders Configuration

Lander Holder CRV Holder



CRV and Lander Holders Analysis



Margin of Safety (MS)

• Surface Crack Propagation• Assumptions:

- Leak before break- a/c = 1.0- a/t = 1.0- a/b = 0.1

Surface Crack Propagation (Fig 8.3 from Fundamentals of Structural Integrity by Alten F.

Grandt)

280 1055.2

mN

t

rpdesign

281028.6,,,

mN

b

c

c

a

t

aM

Q

aK allowfallow

%1471design

allowMS

K = 36.26 MPa-m1/2 for Al 2219-T851



Hoytube Design

• 6 Hoytubes within the bundle

• 5 primary lines per Hoytube

– Most of load bearing capability

• 8 Secondary lines per Hoytube

– Initially slack, load bearing in case of damaged primary lines

• High survivability

– 100 % > 70 years



Component Masses

Hab Component Mass (kg)

Stringers 360

Rings/Frames 620

Columns 2,750

Braces 1,180

Floor Partitions 15,570

Hab Component Mass (kg)

End Caps 4,500

Outer Shell 1,880

Inner Wall 3,980

Micrometeorite Protection 2,500

Airlocks 3,260



Layering System

Layering Thickness

Al 6061 Bumper 2 mm

MLI 6.4 mm

Polyethylene 7 cm

Al 2219-T851 Shell 2 mm



Rigid Body Model

9.31 m Stringer

Ring/Frame

4.38 m



Power Subsystems Breakdown (Primary Power)

Subsystem Power Allotted (kWe)

Percentage of Total Power

Human Factors 23 11.5%

Thermal 20 10%

Power 5 2.5%

Communications 20 10%

Propulsion 0.8 0.4%

Structures 1.5 0.8%

Aerodynamics 0 0%

Margin 9.7 4.8%

Continuous Power Subtotal 80 40%

Dynamics and Controls 120 60%

Total 200 100%

Ryan Spalding, Reuben Schuff, Justin Tucker


Power Cable



Power Cable Mass

Mass Breakdown for Components of Power Cable Copper Silicone Shielding

Density [kg/m^3] 8920 1150 9000

Resistivity [Ohm-m] 1.70E-08 NA NA

Length [m] 75 75 75

Cross-section Area [mm^2] 14 19.2 25.4

Mass per Wire [kg] 9.4 1.7 NA

Number of Wires per Bundle 4 4 NA

Mass per Bundle [kg] 37.4 6.6 17.2

Number of Bundles NA NA 3

Total Mass of Cable [kg] 183.7



Fuel Cell System Mass

Mass Breakdown of Fuel Cell SystemItem Mass (kg) Volume (m3)

Each Fuel Cell 118 0.16

Fuel Cells (5) 590 0.8

Each LOX Tank 115 0.53

LOX Tanks (2) 230 1.06

Each LH2 Tank 131 1.09

LH2 Tanks (2) 262 2.18

Total Hardware 1082 4

LOX Fuel 893 0.77

LH2 Fuel 112 1.59

Total Fuel 1005 2.36

Total System 2100 4Ryan Spalding, Reuben Schuff, Justin Tucker


Power Subsystems Breakdown (Secondary Power)

Location of Power Use Power Supplied (kWe)

Tether Winch 7.5

Human Factors Considerations 10

Communication/Navigation 7

Thermal Concerns 5

Margin 1

Total (without winch) 23

Total (with winch) 30.5


Location of Power Use Power Supplied (kWe)

Human Factors Considerations 4

Communication/Navigation 4

Thermal Concerns 5

Margin 1

Total 14


Breakdown of Fuel Cell System(Duration and Power Supplied)


Interval Time (hr) Power Supplied (kWe)

Power Capacity (kWe-hr)

Tether Deployment and Retraction:

1 Procedure 3 30.5 91.5

Total Required: 6 18 30.5 549

Margin: 14 42 30.5 1281

Total: 20 60 30.5 1830

Main Engine Burn: First Burn 1.5 23 34.5

Second Burn 1 23 23

Third Burn 0.5 23 11.5

Margin 2 23 46

Total 5 23 115

Aerocapture: 2 3 6

Total n/a n/a 1950


Mass Breakdown of Power Distribution System


Components:

Plasma Contactors (Ground) 159

Transformers:

Large 670

Small Scale 5

Regulators, Converters, charge controllers,etc 1037

TOTAL COMPONENTS 1872

TOTAL WIRING 3461

TOTAL DISTRIBUTION SYSTEM 5330


Power Loss in Tether

AreaSurfaceWire

LossPowerqqq emittedsolarnet

Energy Balance at Outter Insulation Surface:

Energy Balance at Inner Insulation Surface:

emittedsolar qqAreaSurfaceWire

LossPower

, Matthew Branson, Lucia Capdevila, Alessandro Ianniello, RobertManning

MelanieSilosky


Cooling Loop Design

• Propulsion Module– Two phase H2O loop

– Mass flow rate = 0.04 kg/s

– Pressure = 2 atm

• Habitat Module– Single phase liquid NH3 loop

– Mass flow rate = 0.08 kg/s

– Supply temperature = 4.4 oC

380 kW

380 kWFromEngines

P

HX

T1 = 130.8 oC

H2O vapor

T2 = 130.8 oC

H2O liquid

HX

HX

P

P

T2 = 4.4 oC

T1 = 83 oC

T2 = 4.4 oC

T1 = 83 oC

33 kW

33 kW


Panel Design

• Panel Design– Beryllium fins (k = 220 W/m-K)

– Z-93 white paint coating ( = 0.92)

10 cm

0.58 mm3.81 mmfin

heatpipe


Radiator Mass Breakdown

Propulsion module Habitat Module

Panel 781 kg 365 kg

Support structure 3780 kg 1763 kg

Total 4561 kg 2128 kg


Timeline

• Early November 2009 – 500 km circular orbit at 23.45º inclination• Late November 2009 – Finite burn for trans-Mars injection, Δv =

4.50 km/s• Mid January 2010 – Tether deployed, spin-up maneuver, ω = 5 rpm• Early June 2010 – Spin-down maneuver, EVA performed, prepare

for aerocapture • Mid June 2010 – Mars atmospheric probes released• Mid July 2010 – Aerocapture into 14 day elliptic orbit around Mars, e

= 0.97• Late July 2010 – First Mars Lander released, landing at 1.98ºS,

353.82ºE• Early August 2010 – Second Mars Lander released, landing at

8.92ºN, 205.21ºW• Mid August 2010 – Apo-twist maneuver• Mid September 2010 – Spin-up maneuver, simulate Mars gravity

Allison Bahnsen, Daniel Grebow, Kelli Hsieh, Steven Lambert, Joseph Paunicka, Brian Pramann


Mars Aerocapture: Capturing the Corridor• Vehicle Characteristics Unchanged

• Entry Corridor Density Uncertainties

• % Cases Captured: 54 Total

Ellipsled

Image taken from R. Whitley and C. Cerimele

Parameter Variation

Standard Dev. -3, 0 and 3

Dust Level Low, Mod, High

Time of Day 0-24 hrs (4 hr incr.)

St. Dev % LU Capt. % LD Capt.

-3 83.33 % 83.33 %

0 100 % 100 %

3 100% 33.33%

Nominal Flight Path Angles [LU, LD] [-9.43º, -8.1065º]

Ryan Whitley


Spin-up/Spin-down Specifics

Spin-Up Spin-down

ΔV (m/s) Time (days) ΔV (m/s) Time (days)

Trans-Mars

Hab side 28.8 41.5 28.8 41.5

Propulsion side 38.2 41.5 38.2 41.5

Mars Orbit

Hab side 10.4 29.5 10.4 28

Propulsion side 13.6 29.5 13.6 28

Trans-Earth

Hab side 28.8 22.1 28.8 20.6

Propulsion side 57.64 22.1 57.64 20.6



Hall Effect Thruster Placement



• Early November 2009 – Initial

Earth parking orbit.

• Late November 2009 – Trans-

Mars injection, 1.34 hour burn

time.

– Impulsive: ΔV = 3.55

km/s.

– Finite: ΔV = 4.50 km/s.

Trans-Earth Injection: Finite Burn

Daniel Grebow, Allison Bahnsen, Kelli Hsieh, Steven Lambert, Joseph Paunicka, Brian Pramann


Finalized Orbital Parameters


TOTALS

Total Mission Time (yrs) 2.36

Total Main Engine ΔV (km/s) 7.88

a(km)

e rp

(km)

ra

(km)

v∞

(km/s)

ΔV(km/s)

P(days)

TOF(days)

Trans-Mars 1.89e8 0.21 1.50e8 2.28e8 2.94 3.55 518 259

HyperbolicArrival

8.44e3 1.43 3.45e3 - 2.64 (ΔVeq = 0.52) - -

Post-CaptureElliptical

1.17e5 0.97 3.45e3 2.30e5 2.67 1.6e-3 13.99 6.99

Mars Parking 1.17e5 0.97 3.60e3 2.30e5 5.01 3.8e-2 14.00 336

“Parabolic”Departure

2.01e8 1.00 3.60e3 - 2e-4 ΔVcr = 2.65 - -

Trans-Earth 1.89e8 0.21 1.50e8 2.28e8 2.93 0.48 518 259


-2 -1.5 -1 -0.5 0

x 105

-8

-6

-4

-2

0

2

4

6

8

x 104

Cartesian X, x [km]C

art

esi

an

Y, y

[km

]

Aerocapture with 14-Day Elliptical Parking Orbit

Hyperbolic ArrivalPost-Capture OrbitElliptical Parking Orbit'Parabolic' Departure

Possible methods to reduce Δvcr:

• Out-of-plane hyperbolic arrival at Mars.• Rotation of the line of apsides and precession of the line of

nodes due to Mars’ oblateness.• Apo-twist maneuvering.• Apply correction maneuver before periapsis.


Mars Elliptical Orbit3,600 km x 230,000 km

v∞

(km/s)

ΔV(km/s)

Trans-Mars Injection 2.94 3.55

Periapsis Raise Maneuver 2.67 0.71

Trans-Earth Injection 5.01 3.8e-2

Correction Maneuver 2e-4 2.65

TOTAL 6.95

Aerocapture into 14-day Elliptic Orbit


Apo-Twist

d/dt=(3*n*J2*Rplanet2(4-5*sin2(i)))/(4*a2(1-e2))

“Squishy”Mars

d

Ecliptic Plane

Equatorial Plane

Orbital Plane

25.19 deg

63.4 deg



-2.5

-2

-1.5

-1

-0.5

0

0.5

1

1.5

2

2.5

x 108

-2

-1

0

1

2

x 108

-20246

x 107

Cartesian X, x [km]

Zubrin's "Athena" Trajectory

Cartesian Y, y [km]

Car

tesi

an Z

, z

[km

]

Earth OrbitMars OrbitInitial Hohmann TransferSpacecraft Intermediate OrbitFinal Hohmann Transfer

Zubrin’s Trajectory



Aerodynamics: Equations of Motion

cos)(

coscos

sincos

tancoscoscos

sin

cos

)()(

cos

)(

sin

sin

2

2

2

hr

V

hr

V

hr

V

mV

L

Vhrhr

V

mV

L

hrm

DV

Vh

mars

mars

mars

marsmars

mars

Ryan Whitley


Aerocapture: Final Altitude Profile

Ryan Whitley


Aerocapture: Final Velocity Profile

Ryan Whitley


Aerocapture: Final G-load Profile

Ryan Whitley


Probe: Equations Used

• Ballistic Coefficient = m/Cd S

• V = Ve exp (1/2Z 1/BC rho/sin gamma (exp –Zh))

• dv/dt = -1/2 1/bc rho V²

• Qrate = k (rho/Rn U/1000)³

Ayu Abdullah


Probe Trajectory (Trajectory found using data from code by Ryan Whitley and Bob Manning)

0 50 100 150 200 250 300 350 400 450-50

0

50

100Altitude versus Time

Time (s)

Alt

itu

de

(k

m)

80 90 100 110 120 130 140 150

2

3

4

5Velocity versus Time

Time (s)

Ve

loc

ity

(k

m/s

)

Ayu Abdullah


Probe Trajectory (Trajectory found using data from code by Ryan Whitley and Bob Manning)

0 100 200 300 400 500 600 700 800-50

0

50

100Altitude versus Range

Range (km)

Alt

itu

de

(k

m)

0 100 200 300 400 500 600 700 8000

2

4

6Velocity versus Range

Range (km)

Ve

loc

ity

(k

m/s

)

Ayu Abdullah


Probe Characteristics

Powered by 2 non-rechargeable lithium-thionyl cloride batteries of 600 miliamp hours, 6 – 14 volts for 1-3 days.

Probes encased in aeroshells made of ceramic material Probes will contain batteries, accelerometers, sun sensor, temperature sensor,

communications equipment. Propulsion system¹

Main engine – Marquadt R6 – C

Two tanks using fuel – N2O4, oxidizer – MMH

3 Retro-rockets which provide Δv = 16 m/s

¹ Propulsion system designed by Nikolaus Ladisch using trajectory from module designed by Brian Pramann

Ayu Abdullah


Visualization of the Mass Breakdown

S1 C1 S2 C2 S3 C3 S4 C4 S5 C5 S6 C6

Matthew Branson, Bob Manning, Alessandro Ianniello, Melanie Silosky, Lucia Capdevila


Ablator Materials4

SLA-561V is a mixture of Silicone, silica microballons, corks and silica glass fibers that is injected into a glass reinforced polymide

honeycomb.

Ablator Materials is used to help cut down on weight. The material is burnt up while entering into an atmosphere to remove some of the heat that is generated while entering.



Composites Purpose in Heat shields

• The main purpose is for Strengthing the heatshield when dealing with such high thermal loads

• Si-C (Silcon-Carbide) is used for its VERY high (2800oK)5 melting point while still maintaning its strength (200-350 MPa)5

• C-C (Carbon-Carbon) is used for its very high (20600K)5



References

1) David G. Gilmore, Spacecraft Thermal Control Handbook, The Aerospace Press, El Segundo, CA., 2002

2) Charles D. Brown, Elements of Spacecraft Design, AIAA Education Series, Castle Rock, CO, 2002

3) Wiley J. Larson and Linda K Pranke, Human Spaceflight, The McGraw-Hill Companies, inc., New York, NY

4) K. Sermeus, Euroavia / Mission to Mars Symposium

5) http://www.ultramet.com/old/therm.htm6) Soddit Matlab code written by Damon Landau7) Sandia One-Dimensional Direct and Inverse Thermal Code (Soddit), Sandia

National Laboratories, Albuquerque, New Mexico, 19908) Professor Schnider


Lander Placement

Ayu Abdullah, Masaaki Atsuta, Allison Bahnsen, Franklin Hankins Leigh Janes, Andy Kacmar, Matt Maier, Dan Nakaima, Ben Phillips


Lander Separation


• Release of First Lander– Correction v =

1.01 m/s

• Release of Second Lander, waiting half a sol:– Correction v =

1.17 m/s

Trajectory of 1st Lander

Transport Trajectory

Trajectory of 2nd Lander


Rover Communication

• Can view half of Mars for 99.73% of the time

• Meets needs of communications– Equatorial Landings Sites are suitable

View from Spacecraft

Equator

-5000 0 5000-8000

-6000

-4000

-2000

0

2000

4000

6000

8000

Cartesian x [km]

Car

tesi

an y

[km

]

Communication Availability By: Allison Bahnsen


Transport OrbitSwath Width


Rover Communication – Swath Width

• Sw = 2**Rs

• At apoapsis, Sw = 10,572 km

• At periapsis, Sw = 2,276 km

• Calculated the distance when Sw = 2*3397km to find when we could see the whole planet

RS

ra


Rover Landing Sites

|203 W | 205W |207 W

10 N_

9 N_

8 N_

8.92° N, 205.21° W


1.98° S, 6.18° W

Athabasca VallesTerra Meridiani


Details on Cruise Stage

5 m

Thrusters Prop. Tank

Sun Sensor

Solar Panels

Star Scanner

Heaters



Parachute SystemVariable Name Material Specific Weight

WC (canopy) Nylon/Kevlar .0115 lb/ft2

WSL (suspension

lines)

Kevlar .0035 lb/ft/1000 lb strength

WRT (radial tape) Kevlar .0035 lb/ft/1000 lb strength

WR (riser) Kevlar .0035 lb/ft/1000 lb strength


Upper portion of Lander and parachute cables

Parameters Drogue Lander

SO [m2] 170 385

DO [m] 10.4 16.7

NSL 48 48

LSL [m] 16 23

NR 1 5

LR [m] 5 3

NG 48 48

Volume [m3] .021 .039

Total mass [kg]

17 32


Aeroshell Ballistic Trajectory

• Entry parameters –

• Ventry = 4.896 km/s

• Gamma = 4.596 degrees

• Ballistic coefficient

• = 99.07 kg/m^2

• Maximum Heating Rate

• = 322.03 W/cm^2

• Altitude of Maximum Heating Rate =35.87 km

• Maximum Deceleration = 4.4 Earth G’s

• Altitude of Maximum Deceleration = 26.31 km

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 50

50

100

150Altitude versus Velocity

Velocity (km/s)

Alti

tude

(km

)

0 0.5 1 1.5 2 2.5 3 3.5 4 4.50

50

100

150Altitude versus Deceleration

Deceleration (Earth Gs)

Alti

tude

(km

)



Aeroshell Design• Cd of Aeroshell =1.69

• Mass of Aeroshell = 435 kg

• -Heatshell = 230 kg - Backshell = 205 kg



Lander Trajectory

0 50 100 150 200 250 300 350 400 450-50

0

50

100Altitude versus Time

Time (s)

Alt

itu

de

(k

m)

0 50 100 150 200 250 300 350 400 4500

1

2

3

4

5Velocity versus Time

Time (s)

Ve

loc

ity

(k

m/s

)



Lander Trajectory

0 200 400 600 800 1000-50

0

50

100Altitude versus Range

Range (km)

Alt

itu

de

(km

)

0 200 400 600 800 10000

1

2

3

4

5Velocity versus Range

Range (km)

Vel

oci

ty (

km/s

)



Aeroshell FEM Analysis


Parameter Maximum value

von Mises stress 2.24 104 N/m2

Displacement 4.62 mm

Compressive stress 2.14 104 N/m2


Graphite Ablation

Carbon-Carbon Composite

Honeycomb

Heat Shield Analysis

Parameter Value

BC 49.07 kg/m2

Maximum G-loading 5.03 Earth G’s

Estimated cross range 727 km


Material of Each Layer Thickness (cm)

Graphite Ablator 0.1

Carbon-Carbon Composite 0.1

Glass Reinforced Polyimide Honeycomb

10


Retro Rocket Specifics

ΔV mfinal tb Pc ε

85 m/s 1575 kg 40 s 3 MPa 30

Isp cF c*

364 s 1.915 1865 m/s


F Dstop Rthroat Rexit Rcham Lcham Lnoz

1739 N 2408 m 0.0098 m 0.054 m 0.0252 m 0.193 m 0.131 m


Retro Rocket Configuration



Lander Dimensions

Height (m)

Thickness (cm)

Mass (kg)

1.1 2 14.3

1.1 2 15.6

N/A 1 44.5

N/A 10 444.9

Total Mass (kg) 609.0


Panel Number Length (m)

Side A 4 1.3

Side B 4 1.4

Top 1 N/A

Bottom 1 N/A

Leg ATop Panel

Leg BBottom Panel

Side Panel A

Leg B

Side Panel B

Leg A

Leg Number Length (m) Diameter (cm) Mass (kg)

A 4 0.95 5 14.7

B 8 1.0 5 15.4

Total Mass (kg) 182.0


Lander Communication

Lander to Rover

Frequency 0.42 GHz

Efficiency Transmitting 0.65

Efficiency Receiving 0.65

Bit Error Rate 5.00e-6 bps

Link Margin 2 dB

Noise Temperature 300 K

Atmospheric Loss 2 dB

Distance of Transmission 1 km

Data Rate 2.00e-4 bps

Power 0.081 mW

Mass 0.0365 kg


Lander to Transport Vehicle

Frequency 21.2 GHz

Diameter Receiving 2 m




Link Margin 2 dB



Distance of Transmission 229,700 km

Data Rate 10 Mbps

Diameter Transmitting 0.32 m

Power 10 W

Mass 0.4 kg


Power Specifics

Lander Power SystemMass of radio-isotope: 60 kgMass of batteries for landing: <1 kgVolume of power systems: 0.2 sq metersPower produced: 300 W (at beginning of life)

Rover Power SystemMass: 24 kgPower Produced: 120 WVolume: ~0.1 sq meters

Failure RateBased on previous missions using radio-isotope power sources the failure rate

for both the lander and rover is <1% (no moving parts)



Rover Specifics

• Mass: approximately 155 kg • Wheelbase (front to rear): 1.2 m • Wheel Size: ~ 0.25 m diameter, 0.15 m width • Track Width: 1.1 m (outside of wheel to outside of

wheel) • Maximum Obstacle Height: 0.30 m rock • Top Deck Height: approx 0.6 m above ground • Rover Body Dimensions: approximately 0.6 x 1.0 x

0.3 m • Mast Instrument Platform Height: 1.0 m above

ground • Arms : 6 degree of freedom (DOF) • One Sol Range: Terrain dependent (50 m Nominal) • Guidance, Navigation & • Control Sensors: Cameras, LN-200 • Effective Stereo Range (Navcams) ~50 m • RPS Power: 200 W continuous (2 RPSs) • Thermal Control: Heat from RPS: Cool from waste

from RPS • Landed Operational Lifetime: 365 Earth Days



Rover Detailed Mass Budget

system component mass / each # mass / all total / system

Mobility System Wheel 2.757 6 16.54 20.956Actuator 0.13 10 1.3Frame 1.558 2 3.116

Arm(L) Arm 0.65 2 1.3 3.168Motor 0.13 6 0.78Gripper 0.6 1 0.6Scoop 0.288 1 0.288Sensor 0.2 1 0.2

Arm(R) Arm 0.65 2 1.3 7.48Motor 0.13 6 0.78Raman Spectrometer 4.3 1 4.3APX 0.8 1 0.8MI 0.3 1 0.3

Head Panacam 0.27 2 0.54 8.31Navcam 0.22 2 0.44Mini-TES 2.1 1 2.1Motor 0.13 4 0.52Mast 4.71 1 4.71

Body Hazcam 0.245 4 0.98 115.214Radiation Detector 5.7 1 5.7Sample Container 0.213 1 0.213HGA 5.7 1 0.867Moror 0.13 2 0.26UHF Antenna 0.034 1 0.034(Motor) 0.13 12 1.56Warm Electronics Box 18 1 18

REM 45.9 1 45.9 IMU 0.7 1 0.7

RPS 40 1 40COMM HW 1 1 1

TOTAL 155.128



Rover Communication

Rover to Lander

Frequency 0.41 GHz




Link Margin 2 dB



Distance of Transmission 1 km

Data Rate 2.00e-4 bps

Power 0.22 mW

Mass 0.0374 kg


Mars Rover to Transport Module

Frequency 21.2 GHz

Diameter Receiving 2 m




Link Margin 2 dB



Distance of Transmission 229,700 km

Data Rate 10 Mbps

Diameter Transmitting 0.32 m

Power 10 W

Mass 0.4 kg


SRV Specifics

Component Component Component

Overall Height 3.02 [m] Take Off Mass 950 [kg] Ispvac 344 [s]

Max Radius 0.48 [m] Dry Mass 200 [kg] Mix Ratio 2.99

Tank Height 2.42 [m] Payload 10 [kg] Chamber P 300 [psi]

Radius 0.48 [m] Fuel 740 [kg] Area Ratio 15

Nozzle Length 0.30 [m] Engines 3 Thrust Coefficient

1.707

Exit Radius 0.11 [m] Thrust/Weight 4.54

Throat Radius 0.03 [m] Total Thrust 16,400 [N] Characteristic Velocity

6064

Cargo Bay Height 0.10 [m] Burn Time 306 [s]

Docking Probe Length

0.20 [m]0.20 [m]

Equivalent V 5.2 [km/s]



Propellant Production Specifics

Methane Oxygen Component

Mass Needed

185 [kg] Mass Needed

550 [kg] Required Hydrogen

47 [kg]

Production Rate

.616 [kg/day] Production Rate

2.46 [kg/day] Production Equipment

20 [kg]

Time 300 [days] Time 223 [days] Power Required

400 [kw]


•Reaction•3CO2 + 6H2 → CH4 + 2CO + 4H2O•2H2O → 2H2 + O2

•1 kg H2 → 3.98 kg Methane & 7.94 kg O2


Launch Parameters


Parameter Numeric Value

Altitude [km] 100

Range [km] 732

X-Velocity [km/s] 4.91

Hohmann speed at 100 km [km/s] 4.91

Burn Time [s] 307

Thrust [N] 13,000


Optimal Launch of SRV

• Two-Point Boundary Value Problem Optimization– Used code created by

Professor Williams

)cos(

)sin(

)sin(

)cos(

2

gm

Tv

m

Tv

vy

vx

y

x

y

x

Initial Conditions Final Conditions

to yf = rc = 100 km

xo vxf = vc = 4.91 km/s

yo vyf = 0

vxo

vyo


Optimal Launch



SRV Docking Views


Fixed End Counter Clockwise rotation of 60°Fixed End Counter Clockwise rotation of 60°


AMCM

Brady Kalb

Cost = αQβMΞδSε(1/(IOC-1900))BφγD

Constants

α = 5.65e-4

β = 0.5941

Ξ = 0.6604

δ = 80.599

ε = 3.8085e-55

φ = -0.3553

γ = 1.5691

Variables

Q = Quantity

M = Dry Mass (kg)

S = Specification

IOC = Initial Operating Capability

B = Block Number

D = Difficulty


AMCM Values

Brady Kalb

Specification IOC Block Number

Difficulty

Launch Vehicle

1.93 2009 2 -1

Transport 2.39 2009 1 0

Lander 2.46 2009 2 -0.5

Rovers 2.14 2009 2 -0.5

Crew Return Vehicle

2.27 2009 3 -1


Cost Schedule

Brady Kalb

Cost Fraction =

A(10F2 – 20F3 + 10F4) + B(10F3 – 20F4 + 10F5) + 5F4 – 4F5

Where F equals fraction of project life complete.

For manned mission, A = 0.32 B = 0.68


Inflation Rates

Brady Kalb

Year Rate (%)

1999 2.21

2000 3.36

2001 2.85

2002 1.58

2003 2.28

spring 2004 aae450: slide 1 introduction brady kalb aae 450 – spring 2004 project homer humans...

Documents

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biogen water recycling

final mass kgpropellant

floor component