High Thrust In-Space Propulsion Technology Development R. Joseph Cassady Aerojet

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A GenCorp Company High Thrust In-Space Propulsion Technology Development R. Joseph Cassady Aerojet 22 March 2011

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High Thrust In-Space Propulsion Technology Development R. Joseph Cassady Aerojet. 22 March 2011. Technology Development Needs a Framework. Critics attack the technology development efforts because they tend to “wander in the desert” - PowerPoint PPT Presentation

Transcript of High Thrust In-Space Propulsion Technology Development R. Joseph Cassady Aerojet

A GenCorp Company

High Thrust In-Space PropulsionTechnology Development

R. Joseph CassadyAerojet

22 March 2011

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Technology Development Needs a Framework

• Critics attack the technology development efforts because they tend to “wander in the desert”

• Lack of a defined destination is cited as a flaw by the critics

• It is important to include ties to examine technologies with a framework that allows their relative merits to be examined in an applied manner – not abstract academic considerations!

• In this same vein, it is important to look for synergies between technologies. This should be a Figure of Merit (FOM)

• Elements that serve as building blocks and that are useful to multiple missions / destinations are also desireable – this is another key FOM

An Example…

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Architecture Study Framework

L2

DestinationsLunar Orbit or L-2

NEOsPhobos

Mars Surface

Launch Propulsion OptionsSDLV (Baseline for Comparison)

HC-ORSC CoreHC-GG CoreH2/O2 Core

Solid/Liquid Booster OptionsLiquid Upper Stage Options

In-Space Propulsion Options Crew Cargo LOX/H2 LOX/H2 LOX/CH4 SEP NTR NTR

ISRU

Launch and In-Space Phases linked by:Total in-space mass and volume requirements

Launch Vehicle/in-space hand-off orbitLaunch Manifest

Commonality opportunities

Mission Phases

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DR=Direct ReturnO=Option

Delivered Mass Requirements for Destinations

Multi-Destination Mission Elements enables affordable approach

Element Concept

Element Name Dry Mass, t Maximum Wet Mass, t

L2/Moon NEO Phobos Mars Surface

Orion Vehicle (Direct Return Option) 13 21

DR

DR

DR

DR

Commercial Crew Vehicle (LEO Option) 9 O O O O

Space Habitat (Crew of 4)

19.5 25

Lunar/NEO Cargo Tug (300kW) 4 15 O

Common LOX/LH2 Earth Departure Stage Earth Return Stage (2x 50klbf engine)

16.5 65

Common LOX/LH2 Destination Insertion Stage (1x 50klbf engine)

8.3 33 O O

Exploration Science (all destinations) 1

Mars Cargo Tug (600kW) 8 58

NTR Stage (60klbs thrust, LH2)

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31 O O

NTR Drop Tank (LH2) 10 40 O O

Mars Landers (x3) Descent & Ascent Vehicle Surface Habitat Exploration Equipment Lander

25 50

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In-Space Propulsion Options

• Only included options which are realistic for next 20 years• Performance metrics were defined from already demonstrated ground testing• Complete Stage Mass models were developed for each technology to use in the Concepts of Operations

Element Propellant Specific Impulse, s

Thrust

Cryogenic Propulsion(1 p.432) LOX/LH2 452 67– 222kN (5-50klbf) descent/ascent thrust was

not yet evaluatedSoft Cryogenic Propulsion LOX/CH4 350Semi-Cryogenic Propulsion LOX/RP1 349Nuclear Thermal Rocket(2 p.25) LH2 900Hall Thruster Systems ( p.11) Xenon or Krypton 3000 40mN/kW or 32mN/kWGridded Ion Thruster Systems Xenon 6000 25mN/kW

[i] Manzella, David, et. al., “Laboratory Model 50 kW Hall Thruster,” NASA TM-2002-211887, September 2002.[ii] Herman, Dan, “NASA’s Evolutionary Xenon Thruster (NEXT) Project Qualification Propellant Throughput Milestone: Performance, Erosion, and Thruster Service Life Prediction After 450 kg,” NASA TM-2010-216816, May 2010.[iii] Aerojet, “NASA Completes Altitude Testing of Aerojet Advanced Liquid Oxygen/Liquid Methane Rocket Engine,” May 4, 2010.[iv] http://www.astronautix.com/engines/rd58.htm, cited: January 17, 2011.

• For each propulsion option we established several CONOPS options to trade• Crew and cargo split, direct return vs. LEO basing, LMO vs. Phobos, how Orion is used, ISRU, etc

• IMLEO was then calculated for each CONOPs

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Example CONOPS: Crew Segment of NEO Mission (Reusable Space Habitat Version)

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Example CONOPS: Crew Segment of Phobos Mission

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Conclusions from Architecture Comparison

• High thrust in-space propulsion options include:– Lox-hydrogen for Earth departure

– Lox-methane for landers and ascent vehicles

– Nuclear thermal rockets for crew transit

• Each of these shows benefits by itself, but can also be employed in a way in an overall architecture that enhances the standalone merits

• Supporting technologies like ISRU (and SEP) provide major combinative benefit

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Final Comment

• Selection of one technology as a principal thrust can have ripple impacts

• From the example:– If ISRU were selected as a key long term investment priority, then a

focus on lox-methane for deep space cryo stages (not EDS) would be advised

– If NTR is selected as a key long term technology, then CFM for long duration storage of hydrogen would be advised and perhaps use of lox-hydrogen for deep space cryo stages is better

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Thank you for the opportunity to present