parametric design -...

Parametric DesignPrinciples of Space Systems Design

Parametric Design

¥ The Design Process¥ Example: Project Diana


AkinÕs Laws of Spacecraft Design - #3

Design is an iterative process. Thenecessary number of iterations is onemore than the number you havecurrently done. This is true at anypoint in time.


Overview of the Design Process

System-level Design(based on discipline-

oriented analysis)

Vehicle-level Estimation(based on a few

parameters from prior art)

System-level Estimation(system parameters based

on prior experience)

Program Objectives ✑System Requirements


Project Diana - Mission Statement

Design a system to return humans tothe moon within the shortest feasibletime span for the minimum achievablecost.


Project Diana - Requirements Document (1)

¥ System must be based on the use ofAmerican launch vehicles in operationalstatus as of 2005

¥ System shall be at least as capable forlunar exploration as the G-mission ApollosystemÐ Two landed crewÐ 24 hour stay time, 3 hour EVAÐ 100 kg of science payload (each way)


American Heavy-Lift Vehicles (2005)

¥ Space Shuttle -25K kg to LEO

¥ Delta IV Heavy -25K kg to LEO


DV Requirements for Lunar MissionsTo:

From:

Low EarthOrbit

LunarTransferOrbit

Low LunarOrbit

LunarDescentOrbit

LunarLanding

Low EarthOrbit

3.089km/sec

LunarTransferOrbit

3.089km/sec

0.808km/sec

3.046km/sec

Low LunarOrbit

0.808km/sec

0.022km/sec

LunarDescentOrbit

0.022km/sec

2.247km/sec

LunarLanding

2.861km/sec

2.062km/sec


Mission Scenario 1

¥ What can be accomplished with a single shuttlepayload (25K kg)?

¥ Assume d=0.1, Isp=320 sec¥ Direct landing

Ð LEO-lunar transfer orbit DV=3.089 km/secÐ Lunar transfer orbit-lunar landing DV=3.046 km/secÐ Lunar surface-earth return orbit DV=2.861 km/secÐ Direct atmospheric entry to landing


Scenario 1 Analysis¥ Trans-lunar injection

¥ Lunar landingÐ r=0.3786Ð mLS=1897 kg

¥ Earth returnÐ r=0.4016Ð mER=572 kg

r eTLI

V

gITLI

sp= =-D

0 3724.

m m r kgTLI o TLI= -( ) =d 6809

This scenario would work for amoderate robotic sample returnmission, but is inadequate for ahuman program.


Mission Scenario 2

¥ Assume a single shuttle payload is used to sizethe lunar descent and ascent elements

¥ Assume d=0.1, Isp=320 sec¥ Direct landing

Ð LEO-lunar transfer orbit DV=3.089 km/secÐ Lunar transfer orbit-lunar landing DV=3.046 km/secÐ Lunar surface-earth return orbit DV=2.861 km/secÐ Direct atmospheric entry to landing


Scenario 2 Analysis¥ Lunar landing

Ð r=0.3786Ð mLS=6965 kg


¥ Trans-lunar injectionÐ r=0.3724

Payload mass still too low forhuman spacecraft. Need a betterfit to launch vehicle constraints.(Ideal situation is to ensure 100%load factors.)

mm

rkgLEO

TLI

TLI

=-( )

=d

91 780,


Mission Scenario 3

¥ Assume three shuttle missions carry identicalboost stages which perform TLI and part ofdescent burn

¥ Fourth shuttle payload completes descent andperforms ascent and earth return

¥ All other factors as in previous scenarios


Scenario 3 Standard Boost Stage

¥ mo=25,000 kg¥ mi =2500 kg¥ mp =22,500 kg¥ LEO departure configuration is three boost

stages with 25,000 kg descent/ascentstage as payload

¥ mLEO =100,000 kg


Scenario 3 TLI Performance¥ Boost stage 1

Ð VTLI remaining=2290 m/sec

¥ Boost stage 2Ð r=0.7; DV2=1119 m/secÐ VTLI remaining=1171 m/sec

¥ Boost stage 3Ð r=0.55; DV3=1875 m/secÐ Residual DV after TLI=704 m/sec

DV gIm m

mm

spLEO prop

LEO1 799= -

-æ

èç

ö

ø÷ =ln sec


Alternate Staging Possibilities

¥ Three identical stages¥ Serial staging (previous chart) DV=3793 m/sec¥ Parallel staging (all three) DV=3525 m/sec¥ Parallel/serial staging (2/1) DV=3750 m/sec¥ Pure serial staging is preferred


Scenario 3 Ascent/Descent Performance

¥ 704 m/sec of lunar descent maneuver performedby boost stage 3

¥ Remaining descent requires 2342 m/secÐ r=0.4738Ð mLS=9346 kg


Return vehicle mass iscomparable to the Geminispacecraft - should be adequatefor adoption as initial baseline.


Baseline System Schematic - Project Diana

Boost Stage 125,000 kg



Descent/LandingStage

15,650 kg

Ascent Stage6528 kg

Crew Cabin2818 kg


Variation 1: Lunar Orbit Staging¥ Assume same process as baseline, but use lunar

parking orbit before/after landing¥ Keep boost stage scenario as before

Ð Residual DV=704 m/secÐ Additional 104 m/sec required for braking into LLO

¥ Add LOI completion to descent stageÐ Descent DV=2269 m/sec + 104 m/sec LOI (r=0.4692)Ð Ascent DV=2084 m/sec (r=0.5145)Ð TEI DV=808 m/sec (r=0.7729)


Variation 1 Results

¥ Lunar landed payload = 9625 kg¥ Payload returned to lunar orbit = 3990 kg¥ Earth return payload = 2685 kg¥ Net loss of 133 kg for crew cabin/entry

vehicle¥ Simplistic approach (includes taking earth

return propellant to/from lunar surface)basically adds another stage, but reducesuseful payload


Variation 2: Lunar Orbit Rendezvous¥ Use specialized vehicle for descent/ascent¥ Keep boost stage scenario as before


¥ Leave TEI stage in lunar orbit duringascent/descentÐ Descent DV=2269 m/sec (r=0.4850)Ð Ascent DV=2084 m/sec (r=0.5145)Ð TEI DV=808 m/sec (r=0.7729)


Variation 2 Mass Calculations

¥ TEI stage inert mass=2500 kg (d=0.1)¥ Propellant for LOI completion=815 kg¥ Desired quantity=crew module mass (mcm)¥ Ascent vehicle mass=2.412 mcm

¥ Descent vehicle mass=2.597 mascent =6.265 mcm

¥ Trans-earth injection maneuverÐ Initial mass=25,000-815-5.265 mcm

Ð Final mass=2500+mcm

Ð r=0.7729 --> mcm=3194 kg (gain of 376 kg = 11%)


Variation 2 System Schematic




Descent Stage12,310 kg

TEI Stage4990 kg

Ascent Stage4510 kg

Crew Cabin3194 kg


Variation 3: Maximal LOR¥ Use specialized vehicle for descent/ascent¥ Keep boost stage scenario as before


¥ Leave earth-return hardware (heat shield,parachutes - assume 1000 kg) and TEI stage inlunar orbit during ascent/descentÐ Descent DV=2269 m/sec (r=0.4850)Ð Ascent DV=2084 m/sec (r=0.5145)Ð TEI DV=808 m/sec (r=0.7729)


LOR Component Definitions

¥ Five components to vehicle at lunar arrivalÐ Boost stage 3 (jettisoned during lunar orbit

insertion)Ð TEI stage (also performs final LOI burn)Ð Descent stageÐ Ascent stageÐ Crew module/entry vehicle, composed of

¥ Earth entry hardware (assumed 1000 kg)¥ Crew module

¥ Arrival mass 25,000 kg


Variation 3 Mass Calculations

¥ TEI stage inert mass=2500 kg (d=0.1)¥ Propellant for LOI completion=815 kg¥ Desired quantity=crew module mass (mcm)¥ Ascent vehicle mass=2.412 mcm

¥ Descent vehicle mass=2.597 mascent =6.266 mcm

¥ Trans-earth injection maneuverÐ Initial mass=25,000-815-5.266 mcm

Ð Final mass=2500+mcm +1000Ð r=0.7729 --> mcm=2997 kg --> mentry =3997 kg


Revised Baseline System Schematic




Descent Stage11,550 kg

TEI Stage5220 kg

Entry Hardware1000 kg

Ascent Stage4232 kg

Crew Cabin2997 kg


The Next Steps From Here

¥ Perform parametric sensitivity analysesÐ Inert mass fractionsÐ Specific impulseÐ Size of entry package left in orbit

¥ Investigate reduction from 3 to 2 boost stagesfor minimal system

¥ Examine growth options allowed by systemÐ One-way cargo variantsÐ Larger module with more boost modules


Overview of the Design Process

System-level Design(based on discipline-

oriented analysis)

Vehicle-level Estimation(based on a few

parameters from prior art)

System-level Estimation(system parameters based

on prior experience)

Program Objectives ✑System Requirements

parametric design -...

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