Transcript SAB Brief

An Overview of Advanced Concepts for
Space Access
Andrew Ketsdever
Marcus Young
Jason Mossman
Anthony Pancotti
44th Joint Propulsion Conference and Exhibit
July 21-23, 2008
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Introduction
•
AFRL Advanced Concepts Group performed critical review of advanced
technologies for space access.
•
Room for improvement?
•
Thrust
Efficiency
Energy
Efficiency
Payload Mass
Fraction
Cost/kg
($1000/kg)
Cost/
Energy Cost
97% (SSME)
0.2
.01
10s
.0001
Technologies Considered:
Using Propellant
Propellantless
•Nuclear
•Electromagnetic (Rail)
•Space Tug
•Elevator
•Beamed Energy
•Space Platforms and Towers
•Advanced Chemical
•Gravity Modification and Breakthrough Physics
•Hypersonic Air Breathing
•Launch Assist
• Analysis performed for advanced concepts (15-50 years) is not sufficiently
accurate for more than semi-qualitative comparisons.
•
Qualitatively consider known missions: microsat to LEO and large comsat to
GEO.
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2
Existing State of the Art
•Advanced launch concept must be more than just a new solution.
•Must yield system level performance improvements over SOA.
Microsat to LEO
Large Comsat to GEO
Orbital Minotaur IV
Boeing Delta IV Heavy
•Reduces microsat launch costs by
reusing Peacekeeper boosters.
•4 stage all solid propellant rocket.
•First flight scheduled for Dec. 2008.
•7 successful Minotaur I flights…
•Developed as part of EELV program.
•Reduce costs by 25%.
•Increase simplicity and reliability.
•Increase standardization.
•Decrease parts count.
•Stage 1: 3 CBCs RS-68 (LH2/LO2).
•Stage 2: 1 RL-10B-2 (LH2/LO2) .
•First flight Nov. 20, 2002.
Performance:
•Thrust: I: 2.2MN, II: 1.2MN, III: .29MN.
•1750kg to LEO.
•Minotaur I ~ $30,000/kg.
Performance:
•Stage 1: Sea Level: 8.673MN @ 410s
•Stage 2: At Altitude: 110kN @ 462s
•22,950 kg to LEO.
•~$10,000/kg.
•“Advanced Concepts” have not aided most recent generation!
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Launch Costs
•Technologically feasible to launch 130,000kg to LEO (Ares V).
•What else is important?
•Isp: Propellant cost represents small fraction of overall…
•Responsiveness: Years/months  Weeks/days?
•Cost($/kg): Limitation on type and amount of payload.
25000
Cost per pound of payload
Athena I
Pegasus
20000
15000
Minotaur I
Taurus
10000
Delta II 7320
Titan IV
Delta II 7920
5000
FALCON
Goal
Delta IV M
Delta IV
Heavy
Atlas V
402
Atlas V 552
0
0
10000
20000
30000
40000
50000
60000
Payload to 100nm, 28.5 deg
•Major focus on reducing launch cost (1/10).
•Improved performance (STS): Not successful.
•Reduced performance (EELV): Not quite successful.
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Other Considerations
•Reliability: Likelihood that launch vehicle will perform as expected
and deliver payload into required orbit.
•Typically 0.91-0.95 (Sauvageau, Allen JPC 1998).
•2/3 due to propulsion elements.
•Upper stages less reliable.
•Increasing would decrease insurance costs, improve RLV competitiveness.
•Availability: Fraction of desired launch dates that can be used.
•Responsiveness: Time from determination of desired launch to actual launch.
•Currently measured in months/years.
•Desert Storm: Sept. 1990  Launch Feb. 1992!
•Ideal to have weeks/days/hours capability.
•Extreme Magnitudes
•SSME: P=6GW dthroat=600cm2  10MW/cm2.
•Saturn V: Height: 116m, Diameter: 10m,
Mass: 6.7 million pounds.
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Propellant: Nuclear
•Nuclear materials have extremely high energy densities.
•Fission: 7 x 1013 J/kg at 100% efficiency.
•Fusion: 6 x 1014 J/kg at 100% efficiency.
•~107 – 108 > chemical
•Benefit practical launch systems?
Nuclear powered
upper stage
History
•Nuclear fission rockets first proposed in the late 1940s.
•Variety of concepts exist with Isp from 800s to > 5000s.
•Typically use hydrogen working gas.
•Nuclear propulsion enabling for large interstellar
missions.
•Launch concepts exist.
•NERVA upper stage.
•Primary concerns: system mass, system cost, allowable
temperatures, socio-political.
•Large size limits applications to large payloads.
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Orion
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Propellant: Nuclear Tug
•Nuclear fission propulsion can enable space tugs.
•Reduce the requirements for launch systems?
•Example: mtug (no payload) of 22,000kg, DV = 4.178km/s.
Where is breakeven?
200000
Finert = 0.1
180000
Finert = 0.3
Payload to GEO (kg)
160000
Finert = 0.5
Finert = 0.7
140000
120000
100000
80000
60000
40000
20000
0
0
500
1000
1500
2000
2500
3000
3500
4000
4500
Specific Impulse (sec)
Significant investments required to reduce specific mass of
nuclear systems.
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Propellant: Laser Beamed Energy
•Chemical Propulsion: energy and ejecta same material (neither fully optimized).
•Beamed Propulsion: energy stored remotely so ejecta could be optimized.
•Lasers and microwaves are both proposed for beamed energy launch.
•Both lasers and microwave sources are under continuous development.
•More emphasis on laser propulsion.
•Laser propulsion was first introduced by Kantrowitz in 1972
1. Heat
Exchange
Laser  heat exchanger  flow
Exotic heat exchangers are required.
2. Plasma
Formation
Form plasma in a nozzle to reach high
operating temperatures.
Have high accuracy pointing requirements.
3. Laser
Ablation
Removal and acceleration of propellant via
laser ablation.
More thrust than PLT, but must carry
propellant.
4. Photon
Pressure
Pressure from photons directly used for
propulsion.
Bae’s PLT has shown 3000x amplification.
Still requires higher powered lasers.
Generation: 1MW  1GW
Laser beamed propulsion will take significant money to develop and deploy and
will only service mSat launches in foreseeable future due to required power levels.
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Propellant: mwave Beamed Energy
Source: Parkin and Culick (2004):
•300 gyrotron sources (140GHz,1MW) 
1000kg to LEO.
•Transmission: Frequency very important.
•Atmospheric Propagation.
•Breakdown.
•Coupling Efficiency.
•Generator Size.
•Coupling
•Plasma Formation
•(Oda et al, 2006) Gas discharge formed at focus of
beam. Plasma absorbs beam energy.
•Heat Exchanger
•(Parkin and Culick) Heat exchanger & hydrogen
propellant yield 1000s, payload mass fraction 5-15%.
•Both laser & microwave beamed energy propulsion systems require significant source
(>1GW) and coupling development to yield viable systems for microsatellite launches.
•Overlap with other source applications.
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Propellant: HEDM
•Performance of chemical rocket is critically dependent on propellant properties.




1000.0
T
I sp 
m
•Problem: High Isp typically low density.
•Goal: Find high Isp, density propellant
•1. Strained ring hydrocarbons.
•2. Polynitrogen
•3. Metallic Hydrogen (216MJ/kg).
10000 R
8000 R
6000 R
900.0
800.0
Specific Impulse (sec)
 mi
DV  I sp g ln
m
 f
700.0
600.0
500.0
400.0
300.0
Theoretical Isp
Gamma = 1.15
P1/P2 = 750
200.0
100.0
0.0
0
5
10
15
20
25
30
Exhaust Molecular Weight
Difficulties
•Molecules containing high potential energy are typically less stable.
•Dramatically more expensive (difficult to manufacture, less alternative uses).
•Require new nozzle materials/techniques.
•Wide range of potential materials yielding both near-term and far-term potential improvements, but
with similar technological challenges: less stable, higher operating temperatures.
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Propellant: Hypersonic Air
Breathing Vehicles
•Oxidizer mass fraction >> payload mass fraction for existing launch systems
(30% vs. 1.2% for STS).
•Can atmospheric oxygen be used instead?
Thrust-to-Weight
•SSME: 73.12
•Scramjet ~ 2
•Alternative technologies show significantly higher Isp, but over a limited range
of Mach number.
•Multi-stage systems are required.
•Parallel systems suffer from volume and mass constraints.
•Combined cycle systems require significant development to integrate
flowpaths.
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Combined Cycle Launch Vehicles
RBCC and TBCC
Rocket Based Combined Cycle (RBCC)
Rocket-ejectorRamjetScramjetRocket
Turbine Based Combined Cycle (TBCC)
TurbojetRamjetScramjetRocket
•Both technologies are under development at the
component/initial integration stages.
•Basic demonstration of scramjets has been shown, but
survivable, reusable vehicles have not.
•Development will probably require decades, but may yield a
revolutionary launch technology.
•Could be viable for both launch scenarios
X-51
X-43A
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Electromagnetic Launch: Railguns
•Multiple proposed EM launch technologies: railgun, coilgun, maglev. Acceleration as a function of track length and launch velocity
•Suffer from similar limitations… Only railguns will be discussed.
100000
Acceleration (g)
10000
10 m
100 m
1 km
10 km
100 km
1000
100
10
1
0
2
4
6
8
10
12
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Technical Challenges
Launch Velocity (km/sec)
•Maintain rail integrity.
•Useful high gee payloads must be developed.
Now: Ei=10MJ,m=3.2kg,Vmuzzle=2.5km/s •Pulsed power system must be developed.
•Aero-thermal loads
64MJ (6MA) System Ready > 2020
Navy
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18
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Direct Launch Requirements
•Vmuzzle > 7.5km/s
•E > 10GJ (35GJ muzzle, 44GJ input for 1250kg)
•L > 1km
•Estimated costs: System cost > $1B, 10,000 launches 
$530/kg.
•Potential for cost savings for microsatellites or
small ruggedized payloads in the very far term.
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Space Elevator
•Cable running from Earth’s surface to orbit.
•Idea originated with Tsiolkovsky in 1895. Ribbon to
•No stored energy required.
Counterweight
•Technical hurdles:
•Require extreme tensile strengths.
•Carbon nanotubes?
Climber
•High power requirements.
•Cost.
•Micrometeoroid/orbital debris impact.
•Weather interactions.
•Atomic oxygen/radiation belts.
Beamed
Power
From Liftport
•Significant economic/technical challenges in the short term.
•Long term possibility…
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Space Platforms and Towers
•Physical structures reaching from the earth’s surface to 100km and above.
•Idea has been around for awhile
•More recently several different configurations have been proposed.
•Solid
•Inflatable
•Electrostatic
•Launching from 100km yields only a small amount of the total required
mechanical energy
•Going from <1km to >100km yields significant technological challenges
•Extreme materials properties.
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•Winds
Circular Orbit Kinetic Energy
Energy/Mass [MJ/kg]
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World’s Tallest
Structure
Potential Energy
Total Mechanical Energy
50
40
30
20
10
0
1
10
100
1000
Altitude [km]
10000
100000
•Energy benefit at 100km is small making the development costs
difficult to justify.
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Burj Dubai
(May 12, 2008: 636m of
818m)
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Gravity Modification and other
Breakthrough Ideas
•Large number of breakthrough physics concepts exist.
•Some are based on unproven physics.
•Modification or complete removal of gravity (reduce Ep).
•Tajmar and Bertolami (J. Prop. Power 2005): “gains in terms of propulsion
would be modest (from these concepts) and lead to no breakthrough”
•Inertial mass modification: increase propellant mass as it is expelled out of
vehicle for increased thrust.
•Gravitational mass modification: lead to direct DV reduction. ~1.4km/s if
m 0. GEO 13km/s  3 km/s.
•Gravitomagnetic fields: Lorentz force analog for gravity. Interact with
Earth’s magnetic field to produce thrust. For most configurations very small
thrust levels are produced.
•Some proven physics yields currently unusable systems.
•Casimir force: force is very small and not applicable for launch.
•Antimatter: convert all mass to energy during annihilation.
•Specific energy density of ~ 9x1016 J/kg. Currently limited in production
rate, cost, and storage. Energy return is ~ 10-10.
•No viable systems based on proven physics.
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Launch Assist: Effects
•Can reviewed concepts provide a fraction of
required DV instead of all of it?
•Consider only first stage launch assist technologies.
•Must provide system level performance benefit.
1. Potential Energy Assist
Launch from higher initial altitude.
LEO: Orbits mostly kinetic energy
100km Space Tower: Added 0.968
MJ/kg (26% potential, 2.9% total).
2. Kinetic Energy Assist
•
•
•
Launch with initial velocity
Need several km/s to be
worthwhile.
Encounter problems with highspeed low altitude flight.
3. DV Loss Assist
•
•
7.5-11km/s
1.0-1.5km/s
Circular Orbit Kinetic Energy
Potential Energy
Total Mechanical Energy
70
60
Energy/Mass [MJ/kg]
•
•
•
DVdesign  DVburnout  DVgravity  DVdrag
GEO
50
40
30
Pegasus
20
Mount
Everest
Near
Space
Dirigible
Space
Tower
LEO
(400km)
10
Launch from higher altitude.
Typically represents several % of
total energy.
0
1
10
100
1000
Altitude [km]
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10000
100000
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Launch Assist: Technologies
1. Air Launch
•
Fixed Wing
•
Balloon
Both feasible only for msat
launch.
Pegasus launcher exists, isn’t
any cheaper, possible other
mission benefits.
2. Electromagnetic Launch
•
Railgun
•
Coilgun
•
Maglev
Both gun technologies
potentially feasible only for msat
launch. Need to increase DE by
> 1000x.
3. Gun Launch
•
Gas Dynamic
•
Light Gas Gun
HARP gun fired 180kg projectile
at 3.6km/s. Next gen could
place 90kg in LEO.
SHARP gun 5kg projectile at
3km/s.
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Conclusions
•Significant room for improvement in launch technology.
•Wide range of concepts proposed and being investigated.
•No obvious winners.
mSat  LEO
Comsat  GEO
Challenges
Nuclear
Mass, Cost, Socio-Political
Space Tug
Significant reduction in specific mass of nuclear system required.
Beamed Energy
Generated power levels. Tracking. Coupling.
HEDM
Stability. Toxicity. Cost. Nozzle Materials.
Hypersonics
Scramjets: thermal load. Rapid combustion. Lifetime. High
thrust-to-weight. Significant atmospheric flight.
Electromagnetic
Power source. Rail integrity. High gee payloads. Rail integrity.
Aerothermal loads.
Elevator
Long defect free nanotubes, atomic oxygen, micrometeoroids,
weather, vibrations.
Platforms
Same as elevator. Must define mission benefit.
Breakthrough
No demonstrated phenomena with sufficient propulsive force.
Launch Assist
High gee payloads. Power sources. Aerothermal.
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Conclusions II
•Significant number of remaining technical challenges.
•Solving any single challenge may not enable complete
systems, but may have broad effects.
•High gee payloads & upper stages.
•High temperature nozzles.
•Very high power instantaneous power levels.
•Lightweight power systems.
•Additional concepts are required!
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Announcing the 2008
Advanced Space Propulsion
Workshop (ASPW 2008)
When: Week of October 6, 2008 (TBD)
Where: Pasadena California
Sponsors: NASA Jet Propulsion Laboratory
& Air Force Research Laboratory (Edwards)
Contact: [email protected]
or [email protected]
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