AAE 450- ERV Propulsion

Stephen Hanna

Preliminary Design Analysis

1/23/01

Outline (no show)

Definitions/ Mission Design Philosophy/Strategies Requirements/ Goals/Limitations Options

Launch Vehicles Available ERV launch

• Technology ERV in transit

• Technology/ attitude control/ etc

Conclusion Further study Questions Comments ‘‘No Show’ slides due to time constraints will not be presentedNo Show’ slides due to time constraints will not be presented Any Questions about ‘No Show’ slides should be asked after Any Questions about ‘No Show’ slides should be asked after

presentation is complete ‘Time Permitting’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: Un- Manned Rating (no show)

Simple designs should be used to full extentRedundancy 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 extentRedundancy 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 V for launch from earth ~7.6 km/s and V 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 Russian Eneriga Class of rockets chosen chosen

AdvantagesAdvantages 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

DisadvantagesDisadvantages US companies do not profit; No longer a patriotic

mission Might delay timeliness of launch, RussianPictures 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: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). 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 handlingControl reflectors and internal safety rodsPrimary and auxiliary coolant loop to

ensure nuclear heat removalEmergency 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 DesignPreliminary 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 (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) Thermo chemistry- c Approximate Pc, Pa, and Pe is typically varied to find ideal Evaluate Me (Mach Numb at Exit) Evaluate expansion ratio Evaluate Isp and verify

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 NTRo 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*tbDelta V = g*ISP*log(Mo/Mf) – Int(Drag/M, tb..0)- g*tb

Propellant Flow Rate:• M(dot)= F/ Isp*gM(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]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 ControlAttitude 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

V Thrust Chamber Sys.- Nozzle, Reactor Containment Vessel, Hardware

for Cooling and Feed System50%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 SkillsPropulsion

A&AE 439 A&AE 590K A&AE 590C

Aerodynamics A&AE 415 – CMARC, Aero CAD A&AE 412 – CFD

Other Surf CAM (can use *.dxf flies with minor editing)

CNC machine

References (no show)

INCOMPLETE LISTINCOMPLETE LIST

1. Buden, D. (EG and G energy Measurements, INC.), Tutorial on Nuclear Thermal Propulsion Safety on Mars, AIAA Paper 92-3696, 1992.

2. Sanders, Jerry B. (NASA Johnson Space Center), Manned Space Nuclear Systems Design Guidelines, AIAA Paper 92-3418, 1992.

3. Sivers, R. K. ,Livingstion, J.M. and Pierce, B.L. (Westinghouse Electric Corp.), NERVA propulsion system design considerations, AIAA Paper 90-1951, 1992.