AAE 450- ERV Propulsion
Stephen Hanna
Preliminary Design Analysis
1/23/01
Outline (no show)
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Definitions/ Mission
Design Philosophy/Strategies
Requirements/ Goals/Limitations
Options
 Launch Vehicles Available
 ERV launch
• Technology
 ERV in transit
• Technology/ attitude control/ etc
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Conclusion
Further study
Questions Comments
‘No Show’ slides due to time constraints will not be presented
Any Questions about ‘No Show’ slides should be asked after
presentation is complete ‘Time Permitting’
Define Mission (no show)
Cost
 Lift payload of ~75 tonnes to LEO
 ~50% of Cost of space systems is propulsion.
 Propulsion costs are derived from WEIGHT
Safety
 Proven technology
Mission Success
 Based on combination of Cost and Safety
Design Philosophy: UnManned Rating (no show)
Simple designs should be used to full extent
Redundancy crucial in all critical systems, where
practical. (critical meaning failure to function
causes irreparable mission failure)
High- quality, tests of actual system
hardware/software should be a primary
requirement
Design Philosophy: Man
Rating (no show)
Simple designs should be used to full extent
Redundancy crucial in all critical systems, where
practical. (critical meaning failure to function
causes loss of life)
High- quality, unmanned tests of actual system
hardware/software should be a primary
requirement
Safety decisions should not be influenced by
cost, manpower, or schedule. Risk assessments
should be preformed to determine impact of
these factors on the system.
Design Philosophy:
Minimum Risk (no show)
Critical systems and components where
redundancy is not practical (I.e. structures,
pressure vessels, fittings, etc.) shall be
designed for min risk using conservative
design specifications (factors of safety,
positive life margins, leak before burst ,
etc.).
Strategies
A heavy lift launch vehicle to limit on- orbit
assembly
A split mission strategy (cargo and crew
fly on separate missions)
Pre-deployed and verified “turn- key”
habitats with ‘in situ’ capabilities
Requirements/Limitations
 Launch a Earth Return Vehicle (~2011) with “in sitiu”
propellant production plant and all necessary supplies
(rover, scientific equipment, food, etc..)
 Approximate payload weight for preliminary design ~75 tonnes
• Payload weight estimated using DRM (~100 tons) and Zub. (~45
tons) as boundaries
• Does not include upper stage weights ~100 tonnes and ~95 tonnes,
DRM and Zub., respectively
 Preliminary DV for launch from earth ~7.6 km/s and DV of ~3.7
km/s to get to MARS
• From Team B Orbits Group initial estimates
• Plus an additional 20% Delta V for gravity and aerodynamic losses
• Plus an additional 5% Delta V for Orbit adjustments due time delays
 Approximate payload dimensions similar to DRM and Zub.’s
Requirements/Limitations
 Use existing launch vehicle technology
(modified)
 Why? Cost of Building a “Large Heavy Lifting Vehicle”
is in BILLIONS of dollars for desired safety factor
 Generic H.L.V. for both HAB and ERV is viable
 Final stage of rocket can be modified for specific payload
weight and to use the “no human factors” plus of the ERV
 Nuclear Thermal Rocket for final stage
 Best weight to thrust
 Estimated Cost ~$1 billion dollars
 Attitude Control that is reliable/ robust
Launch Vehicles
Available(no show)
 ALS- NLS
 US Heavy launch Vehicle program
 approximate cost to continue development ~$12 billion (1992
dollars) not including infrastructure and timeline requirements
 Saturn V
 US Heavy launch vehicle (not viable, development estimates
over ~$15 billion)
 Derivatives
 Shuttle SRMB and main engine combination
 Proven technology, unknown costs for heavy modification
 Eneriga Class of Rockets
 Russian, Heavy Launch Vehicle
 Viable option (Built and on the Shelf)
ERV launch(from earth) ]
Best Viable Option: [In conjunction with HAB team requirements]
Russian Eneriga Class of rockets chosen
 Advantages
 System all ready tested and production
facilities and infrastructure present
• To Reduce the cost of development
o Use existing Vehicle, with modifications as
needed
• To reduce the cost of installation of
infrastructure
o Use existing vehicle with infrastructure present
 Modular Design
 World Collaboration spread risk
 Bonus: Rubble, more bang for your buck
 Disadvantages
 US companies do not profit; No longer a patriotic
mission
 Might delay timeliness of launch, Russian
Pictures from http://www.friendspartners.org/~mwade/lvsenergia.htm
Nuclear Thermal Rocket:
Existing Derived Technology for Upper Stage
Three most researched/ funded NTR’s to date:
 NERVA (Nuclear Engine for Rocket Vehicle Technology)
 Flight engine developed in 1970’s
 Lowest Performance NTR in ISP terms
 Most Money invested
 Similar Program  CIS (commonwealth of independent states)
technology is similar but less infrastructure and testing available
 PBR (Particle Bed Technology)
 Highest Performance NTR in ISP terms for solid core design
(high safety factors)
 Similar Program  PBR is NERVA derived and has some technology
crossover that can be applied to NERVA
 Project Conclusion in 1993
 CERMET – Core nuclear rocket
 Good design if reusability an issue, cannot meet timeline
requirements in a costly manner
NTR (NERVA derived):
Record (no show)
 Largest Endeavor for NTR research over twenty years research with
the most money appropriated (1947 USAF Start,1958 NASA runs, & Fin.1972).
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All research available and infrastructure still present for program to be
reactivated
 Flight model complete before project completion and was tested
 28 full power tests with restarts
 Restart time from 1hr to 1 year
 Life span of three years
 Up to 30- minute test durations
 Most time tested on engines 4.5 hrs total
 Reactor sizes ranging from 300 Mw to 200,000Mw
 Development of way to contain affluent (propellant exhaust gases) for
safe ground testing (used on later tests)
 Development of various high temperature fuels
 Solution to problem of fuel erosion of Hot- Hydrogen
NTR (NERVA derived):
Record cont… (no show)
 Specific Impulse as high as 835 sec (avg. 825)
 Recent studies point to improvements saying ISP levels of
~1000 sec
 Thrust levels as high as 890,000 N
 Recent studies sight major possible improvements
 “dual mode”- can be used for onboard electric power and
refrigeration
 Production tested of up to 100Mw
 Proven Ground Technology
 Excellent safety record
NTR (NERVA derived)Reactor Safety (no show)
Neutron poisoning for ground handling
Control reflectors and internal safety rods
Primary and auxiliary coolant loop to
ensure nuclear heat removal
Emergency mode on the order of 30,000
lb- thrust, 500 sec specific impulse, and
10^8 lb-sec total impulse
Assured shutdown at end of life
ERV- NTR(to Mars)
Preliminary Design
 Propellant Chemistry
 NERVA restricted with 70’s technology to 2361K and rods are at
2650K
• Note: A thermal gradient from periphery of the coolant channel wall
(I.e. increasing coolant (propellant) temperature rises resulting in
perphial core temperature increase, melt down)
 Gas properties looked at
• Hydrogen (H2) 2.016 grams/mole
• Methane (CH4) 16.043 grams/mole
• Carbon Dioxide (CO2)  44.01 grams/mole
 Using this isentropic parameter (g) and Cp(T)  C* can be
found
 Results need further interpretation
ERV- NTR(to Mars)
Trade studies in nozzle's expansion ratio to find
ISP (iterative process)
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Thermo chemistry- g, M, Tc
Approximate Pc, Pa, and l
Pe is typically varied to find ideal
Evaluate Me (Mach Numb at Exit)
Evaluate expansion ratio
Evaluate Isp and verify l
ERV- NTR(to Mars)
 Size overall system using Rocket Equation
 Inert- Mass fraction
• Heavier than liquid rockets but have a better thrust to weight
ratio.
• Most systems similar to liquid engines except for radiation
shielding, can double weight of the NTR
o If Hydrogen used on the order of 0.5 to 0.7 using NERVA
research assumed due to lack of extensive database can be
used for initial studies and iterated back to at a later date
 Rocket Equation:
Delta V = g*ISP*log(Mo/Mf) – Int(Drag/M, tb..0)- g*tb
 Propellant Flow Rate:
• M(dot)= F/ Isp*g
ERV- NTR(to Mars)
 Propellant Heating Requirements (important mile
marker)
 From ISP and desired Thrust- mass flow rate can be found
giving us heat needed
 Heat is directly proportional for the power required by the reactor
Pcore=M(dot)*{Hv+int(Cp dt, t2..t1)}= M(dot)*P [Watts]
Example: 1kg/s flow of H2* P=0.018061, T1= 2000K, and T2=3500k
{note a linear correlation} **Assuming gas is initially a liquid
 Pressure drops
 20-30% pressure loss in regenerative cooling, also for making prop a
gas
 Core pressure drops show ~30% but improvements taught ~10%
ERV in Transit
(to Mars)
Attitude Control
 Primary System: Cold Thrusters from
NTR are possible
 Secondary system using MMH and N204
possible (used by STS)
 Further Study Needed
Conclusion (no show)
H.L.V. use present technology, Energia
Class, with modification of final stage as a
NTR propulsion scheme based on NERVA
technology
Further Study
 Finalize Code to size NERVA derived technology based
on D V
 Thrust Chamber Sys.- Nozzle, Reactor Containment Vessel,
Hardware for Cooling and Feed System50%mass
 Propellant Storage System 10%margin of safety
 Pressurant System isentropic blowdown system
 Radiation Shield shield radius equivalent to core radius
 Support Structure 10% of total structure
 More information on NTR use on Mars as a power
source and for return trip using NTR and super heated
CH4
 More in depth look at micro thrusters for ERV attitude
control
 More Cost/ Safety analysis
Question or Comments?
Pertinent Skills
Other
Propulsion
 A&AE 439
 A&AE 590K
 A&AE 590C
 Surf CAM
(can use *.dxf flies with minor editing)
 CNC machine
Aerodynamics
 A&AE 415 – CMARC, Aero CAD
 A&AE 412 – CFD
References (no show)
INCOMPLETE LIST
1.
2.
3.
Buden, D. (EG and G energy Measurements, INC.), Tutorial
on Nuclear Thermal Propulsion Safety on Mars, AIAA Paper
92-3696, 1992.
Sanders, Jerry B. (NASA Johnson Space Center), Manned
Space Nuclear Systems Design Guidelines, AIAA Paper 923418, 1992.
Sivers, R. K. ,Livingstion, J.M. and Pierce, B.L.
(Westinghouse Electric Corp.), NERVA propulsion system
design considerations, AIAA Paper 90-1951, 1992.