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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS1996

High Leverage Space Transportation System Technologies for Human Exploration Missions to the Moon and Beyond

The feasibility of returning humans to the Moon by 2004, the 35th anniversary of the Apollo 11 landing, is examined assuming the use of existing launch vehicles (the Space Shuttle and Titan 4B), a near term, advanced technology space transportation system, and extraterrestrial propellant--specifically 'lunar-derived' liquid oxygen or LUNOX. The lunar transportation system (LTS) elements consist of an expendable, nuclear thermal rocket (NTR)-powered translunar injection (TLI) stage and a combination lunar lander/Earth return vehicle (LERV) using cryogenic liquid oxygen and hydrogen (LOX/LH2) chemical propulsion. The 'wet' LERV, carrying a crew of 2, is configured to fit within the Shuttle orbiter cargo bay and requires only modest assembly in low Earth orbit. After Earth orbit rendezvous and docking of the LERV with the Titan 4B-launched NTR TLI stage, the initial mass in low Earth orbit (IMLEO) is approx. 40 t. To maximize mission performance at minimum mass, the LERV carries no return LOX but uses approx. 7 t of LUNOX to 'reoxidize' itself for a 'direct return' flight to Earth followed by an 'Apollo-style' capsule recovery. Without LUNOX, mission capability is constrained and the total LTS mass approaches the combined Shuttle-Titan 4B IMLEO limit of approx. 45 t even with enhanced NTR and chemical engine performance. Key technologies are discussed, lunar mission scenarios described, and LTS vehicle designs and characteristics are presented. Mission versatility provided by using a small 'all LH2' NTR engine or a 'LOX-augmented' derivative, either individually or in clusters, for outer planet robotic orbiter, small Mars cargo, lunar 'commuter', and human Mars exploration class missions is also briefly discussed.

Borowski, Stanley K.↗
NASA NTRS2008

NASA Experience with Pogo in Human Spaceflight Vehicles

An overview of more than 45 years of NASA human spaceflight experience is presented with respect to the thrust axis vibration response of liquid fueled rockets known as pogo. A coupled structure and propulsion system instability, pogo can result in the impairment of the astronaut crew, an unplanned engine shutdown, loss of mission, or structural failure. The NASA history begins with the Gemini Program and adaptation of the USAF Titan II ballistic missile as a spacecraft launch vehicle. It continues with the pogo experienced on several Apollo-Saturn flights in both the first and second stages of flight. The defining moment for NASA s subsequent treatment of pogo occurred with the near failure of the second stage on the ascent of the Apollo 13 mission. Since that time NASA has had a strict "no pogo" philosophy that was applied to the development of the Space Shuttle. The "no pogo" philosophy lead to the first vehicle designed to be pogo-free from the beginning and the first development of an engine with an integral pogo suppression system. Now, more than 30 years later, NASA is developing two new launch vehicles, the Ares I crew launch vehicle propelling the Orion crew excursion vehicle, and the Ares V cargo launch vehicle. A new generation of engineers must again exercise NASA s system engineering method for pogo mitigation during design, development and verification.

Larsen, Curtis E.↗
NASA NTRS1961

Ignition of Hydrogen-Oxygen Rocket Combustor with Chlorine Trifluoride and Triethylaluminum

Ignition of a nominal-125-pound-thrust cold (2000 R) gaseous-hydrogen - liquid-oxygen rocket combustor with chlorine trifluoride (hypergolic with hydrogen) and triethylaluminum (hypergolic with oxygen) resulted in consistently smooth starting transients for a wide range of combustor operating conditions. The combustor exhaust nozzle discharged into air at ambient conditions. Each starting transient consisted of the following sequence of events: injection of the lead main propellant, injection of the igniter chemical, ignition of these two chemicals, injection of the second main propellant, ignition of the two main propellants, increase in chamber pressure to its terminal value, and cutoff of igniter-chemical flow. Smooth ignition was obtained with an ignition delay of less than 100 milliseconds for the reaction of the lead propellant with the igniter chemical using approximately 0.5 cubic inch (0-038 lb) of chlorine trifluoride or 1.0 cubic inch (0-031 lb) of triethylaluminum. These quantities of igniter chemical were sufficient to ignite a 20-percent-fuel hydrogen-oxygen mixture with a delay time of less than 15 milliseconds. Test results indicated that a simple, light weight chemical ignition system for hydrogen-oxygen rocket engines may be possible.

Gregory, John W.↗
NASA NTRS1994

Metallized gelled propellants: Oxygen/RP-1/aluminum combustion experiments

A series of combustion experiments using metallized gelled liquid propellants were conducted. These experiments used a small 30- to 40-lb(sub f) thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-wt percent, 5-wt percent, and 44-wt percent loadings of aluminum and gaseous oxygen and the oxidizer. Ten different injectors were used during the testing: 6 were for the baseline O2/RP-1 tests and 4 for the gelled fuels. Relatively high C-star efficiencies were obtained with gelled RP-1 (0-wt% RP-1/Al) and metallized 5-wt% RP-1/Al over the O/F range tested: 90-98%. A peak of 98 percent efficiency was reached with ungelled O2/RP-1 and up to 95% efficiency was obtained with gelled RP-1/Al (55-wt% Al). Injector erosion was evident with the 55-wt% testing, while there was little or no erosion seen with the gelled RP-1 with 0 and 5-wt% Al. A protective layer of gelled fuel formed in the firings that minimized the damage to the rocket injector face. This effect may provide a useful technique for engine cooling.

Palaszewski, Bryan↗
NASA NTRS1951

Theoretical performance of lithium and fluorine as a rocket propellant

Theoretical performance for liquid lithium and liquid fluorine as a rocket propellant was calculated with assumptions both of equilibrium and frozen composition during expansion. Parameters included were specific impulse, combustion-chamber temperature, nozzle-exit temperature, composition, mean molecular weight, characteristic velocity, coefficient of thrust, and ratio of nozzle-exit area to throat area. For chamber pressure of 300 pounds per square inch absolute and expansion to 1 atmosphere, the maximum equilibrium specific impulse calculated was 335.5 pound-seconds per pound. The effect of ionization on calculated performance was shown to be negligible by comparison of values of various parameters calculated both with and without ionized products of combustion.

FUELS -ROCKETS (INCLUDES FUELS AND OXIDANT)↗
NASA NTRS1988

Apollo Lightcraft Project

This second year of the NASA/USRA-sponsored Advanced Aeronautical Design effort focused on systems integration and analysis of the Apollo Lightcraft. This beam-powered, single-stage-to-orbit vehicle is envisioned as the shuttlecraft of the 21st century. The five person vehicle was inspired largely by the Apollo Command Module, then reconfigured to include a new front seat with dual cockpit controls for the pilot and co-pilot, while still retaining the 3-abreast crew accommodations in the rear seat. The gross liftoff mass is 5550 kg, of which 500 kg is the payload and 300 kg is the LH2 propellant. The round trip cost to orbit is projected to be three orders of magnitude lower than the current space shuttle orbiter. The advanced laser-driven 5-speed combined-cycle engine has shiftpoints at Mach 1, 5, 11 and 25+. The Apollo Lightcraft can climb into low Earth orbit in three minutes, or fly to any spot on the globe in less than 45 minutes. Detailed investigations of the Apollo Lightcraft Project this second year further evolved the propulsion system design, while focusing on the following areas: (1) man/machine interface; (2) flight control systems; (3) power beaming system architecture; (4) re-entry aerodynamics; (5) shroud structural dynamics; and (6) optimal trajectory analysis. The principal new findings are documented. Advanced design efforts for the next academic year (1988/1989) will center on a one meter+ diameter spacecraft: the Lightcraft Technology Demonstrator (LTD). Detailed engineering design and analyses, as well as critical proof-of-concept experiments, will be carried out on this small, near-term machine. As presently conceived, the LTD could be constructed using state of the art components derived from existing liquid chemical rocket engine technology, advanced composite materials, and high power laser optics.

Myrabo, Leik N.↗
NASA NTRS2009

Gauging Systems Monitor Cryogenic Liquids

Rocket fuel needs to stay cool - super cool, in fact. The ability to store gas propellants like liquid hydrogen and oxygen at cryogenic temperatures (below -243 F) is crucial for space missions in order to reduce their volumes and allow their storage in smaller (and therefore, less costly) tanks. The Agency has used these cryogenic fluids for vehicle propellants, reactants, and life support systems since 1962 with the Centaur upper stage rocket, which was powered with liquid oxygen and liquid hydrogen. During proposed long-duration missions, super-cooled fluids will also be used in space power systems, spaceports, and lunar habitation systems. In the next generation of launch vehicles, gaseous propellants will be cooled to and stored for extended periods at even colder temperatures than currently employed via a process called densification. Densification sub-cools liquids to temperatures even closer to absolute zero (-459 F), increasing the fluid s density and shrinking its volume beyond common cryogenics. Sub-cooling cryogenic liquid hydrogen, for instance, from 20 K (-423 F) to 15 K (-432.4 F) reduces its mass by 10 percent. These densified liquid gases can provide more cost savings from reduced payload volume. In order to benefit from this cost savings, the Agency is working with private industry to prevent evaporation, leakage, and other inadvertent loss of liquids and gases in payloads - requiring new cryogenic systems to prevent 98 percent (or more) of boil-off loss. Boil-off occurs when cryogenic or densified liquids evaporate, and is a concern during launch pad holds. Accurate sensing of propellants aboard space vehicles is also critical for proper engine shutdown and re-ignition after launch, and zero boil-off fuel systems are also in development for the Altair lunar lander.

Source record↗
NASA NTRS2009

Aerojet - Attitude Control Engines

All the engines were both qualification and acceptance tested at Marquardt s facilities. After we won the Apollo Program contract, we went off and built two vacuum test facilities, which simulated altitude continuous firing for as long as we wanted to run an engine. They would run days and days with the same capability we had on steam ejection. We did all of the testing in both for the qualification and the acceptance test. One of them was a large ball, which was an eighteen-foot diameter sphere, evacuated again with a big steam ejector system that could be used for system testing; that s where we did the Lunar Excursion Module testing. We put the whole cluster in there and tested the entire cluster at the simulated altitude conditions. The lowest altitude we tested at - typically an acceptance test - was 105,000 feet simulated altitude. The big ball - because people were interested in what they called goop formation, which is an unburned hydrazine product migrating to cold surfaces on different parts of spacecraft - was built to address those kinds of issues. We ran long-life tests in a simulated space environment with the entire inside of the test cell around the test article, liquid nitrogen cooled, so it could act as getter for any of the exhaust products. That particular facility could pull down to about 350,000 feet (atmosphere) equivalent altitude, which was pushing pretty close to the thermodynamic triple point of the MMH. It was a good test facility. Those facilities are no longer there. When the guys at Marquardt sold the company to what eventually became part of Aerojet, all those test facilities were cut off at the roots. I think they have a movie studio there at this point. That part of it is truly not recoverable, but it did some excellent high-altitude, space-equivalent testing at the time. Surprisingly, we had very few problems while testing in the San Fernando Valley. In the early 1960s, nobody had ever seen dinitrogen tetroxide (N2O4), so that wasn't too big a deal. We really did only make small, red clouds. In all the hundreds of thousands of tests and probably well over one million firings that I was around that place for, in all that thirty-something years, we had a total of one serious injury associated with rocket engine testing and propellants. Because we were trying to figure out what propellants would really be good, we tried all of the fun stuff like the carbon tetrafluoride, chlorine pentafluoride, and pure fluorine. The materials knowledge wasn't all that great at the time. On one test, the fluorine we had didn't react well with the copper they were using for tubing, and it managed to cause another unscheduled disassembly of the facility. It was very serious. It's like one of those Korean War stories. The technician happened to be walking past the test facility when it decided to blow itself up. A piece of copper tubing pierced one cheek and came out the other. That was the only serious accident in all of the engines handled in all those years. Now, we did have a problem with the EPA later because they figured out what the brown clouds were about. We built a whole bunch of exhaust mitigation scrubbers to take care of engine testing in the daytime. In general, we operated the big shuttle (RCS) engine, the 870- pounder at nominal conditions; they scrubbed the effluents pretty well. If you operated that same 870-pound force engine at a level where you get a lot of excess oxidizer, yeah, there s a brown cloud. But, you know, it doesn't show up well in the dark. They did do some of that. But, that s gone; it was addressed one way or another. RELEASED -

Pfeifer, Gerald R.↗
NASA NTRS1989

Laser powered interorbital vehicle

A preliminary design of a low-thrust Laser Powered Interorbital Vehicle (LPIV) intended for cargo transportation between an Earth space station and a lunar base is presented. The selected mission utilizes a spiral trajectory, characteristic of a low-thrust spacecraft, requiring eight days for a lunar rendezvous and an additional nine days for return. The ship's configuration consists primarily of an optical train, two hydrogen plasma engines, a 37.1 m box-beam truss, a payload module, and propellant tanks. The total mass of the vehicle, fully loaded, is 63,300 kg. A single plasma, regeneratively cooled engine design is incorporated into the two 500 N engines. These are connected to the spacecraft by turntables that allow the vehicle to thrust tangential to the flight path. Proper collection and transmission of the laser beam to the thrust chambers is provided through the optical train. This system consists of a 23-m-diameter primary mirror, a convex parabolic secondary mirror, a beam splitter, and two concave parabolic tertiary mirrors. The payload bay is capable of carrying 18,000 kg of cargo and is located opposite the primary mirror on the main truss. Fuel tanks carrying a maximum of 35,000 kg of liquid hydrogen are fastened to tracks that allow the tanks to be moved perpendicular to the main truss. This capability is required to prevent the center of mass from moving out of the thrust vector line. The laser beam is located and tracked by means of an acquisition, pointing, and tracking system that can be locked onto the space-based laser station. Correct orientation of the spacecraft with the laser beam is maintained by control moment gyros and reaction control rockets. In addition, an aerobrake configuration was designed to provide the option of using the atmospheric drag in place of propulsion for a return trajectory.

Clarke, M. T.↗
NASA NTRS2001

Numerical Simulation of One- and Two-Phase Flows in Propulsion Systems

In this report, we present some results of problems investigated during joint research between the Hampton University Fluid Mechanics and Acoustics Laboratory (HU/FM&AL), NASA GRC, and the LaRC Hyper-X Program. This work is supported by joint research between the NASA GRC and the Institute of Mechanics at Moscow State University (IM/MSU) in Russia under a CRDF grant. The main areas of current scientific interest of the HU/FM&AL include an investigation of the proposed and patented advanced methods for aircraft engine thrust and noise benefits. These methods are based on nontraditional 3D corrugated and composite nozzle, inlet, propeller and screw designs such as a Bluebell and Telescope nozzle, Mobius-shaped screw, etc. This is the main subject of our other projects, of which one is presented at the current conference. Here we analyze additional methods for exhaust jet noise reduction without essential thrust loss and even with thrust augmentation. Such additional approaches are: (1) to add some solid, fluid, or gas mass at discrete locations to the main supersonic gas stream to minimize the negative influence of strong shock waves formed in propulsion systems. This mass addition may be accompanied by heat addition to the main stream as a result of the fuel combustion or by cooling of this stream as a result of the liquid mass evaporation and boiling; (2) Use of porous or permeable nozzles and additional shells at the nozzle exit for preliminary cooling of the hot jet exhaust and pressure compensation for off-design conditions (so-called continuous ejector with small mass flow rate); and (3) to propose and analyze new effective methods of fuel injection into the flow stream in air-breathing engines. The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines. The FM&AL Team uses analytical methods, numerical simulations, and experimental tests at the Hampton University campus, NASA, and IM/MSU. The main outcomes during this reporting period are: (1) Publications: The AIAA Paper #01-1893 has been accepted for the AIAA/NAL-NASDA-ISAS 10th International Space Planes and Hypersonic Systems and Technologies Conference, 24-27 April 2001, Kyoto, Japan. The AIAA Paper #01-3204 has been accepted for presentation at the 37th AIAA/ASME/SAE/ASEE Joint Propulsion Conference being held on 08-11 July, in Salt Lake City, UT; (2) Grants and proposals: The HU/FM&AL was awarded the NASA grant NAG3-2495 in October 2000 and the CRDF award was granted to the NASA GRC-HU FM&AL and IM/MSU (Russia) in July 2000. A solicited proposal was submitted for the NASA NRA-01GRC-02 competition and two unsolicited proposals to NASA are in preparation; (3) Theory and numerical simulations: Analytical theory, numerical simulation, comparison of theoretical with experimental results, and modification of theoretical approaches, models, grids, etc. Such investigations have been conducted for three main problems: (a) Combustion efficiency optimization in the half-duct combustor system; (b) Drag reduction effects for blunt bodies with solid needles; and (c) Solid particle injection from the butt-end against a supersonic flow. The NASA CFL3D, HU/FM&AL, and IM/MSU GODUNOV codes were used and modified for solution of these problems. The codes are based on full Euler and Navier-Stokes solvers with and without nonequilibrium oxygen-nitrogen and air-hydrogen chemical reactions in laminar and turbulent gas flow regimes; (4) Experimental tests: In the IM/MSU supersonic wind tunnel, experimental tests were conducted with different number, location, and geometric parameters of solid needles mounted at the front of the butt-end in supersonic flow. Optimal parameters were determined that provide minimal butt-end drag in the stationary flow regime. Experimental and numerical simulation results are in good agreement; (5) Student Research Activity: Involvement of one graduate and two undergraduate students as research assistants in the current project.

Gilinsky, Mikhail↗
NASA NTRS2019

Progress in Additively Manufactured Copper-Alloy GRCop-84, GRCop-42, and Bimetallic Combustion Chambers for Liquid Rocket Engines

Additive Manufacturing (AM) has significantly evolved over the last decade for use in the aerospace industry, particularly for liquid rocket engines. AM offers a considerable departure from traditional manufacturing to rapidly fabricate components with complex internal features. High performance liquid rocket engine combustion chambers that operate in a high heat flux environment are fabricated using a copper-alloy liner with a series of integral coolant channels. Copper-alloys provide the necessary conductivity and material strength for adequate design margins offering high performance without the need for film coolant. Copper-alloys present unique challenges to properly melt the powder in laser-based AM processes due to their high reflectivity and conductivity. Starting in 2014, NASA’s Marshall Space Flight Center (MSFC) and Glenn Research Center (GRC) have developed a process for AM of GRCop (Copper- Chrome-Niobium) alloys using Selective Laser Melting (SLM). GRCop, originally developed at GRC, is a high conductivity, high-strength, dispersion strengthened copper-alloy for use in high-temperature, high heat flux applications. NASA has completed significant material characterization and testing, along with hot-fire testing, to demonstrate that GRCop-42 and GRCop-84 alloys are suitable for use in combustion chambers. Additional development and testing has been completed on AM bimetallic chambers using GRCop-84 liners and superalloy jackets, fabricated using two Directed Energy Deposition (DED) processes: Electron Beam Freeform Fabrication (EBF³) and Blown Powder DED. NASA completed hot-fire testing on various AM chambers using GRCop-84, GRCop-42, and bimetallic chambers in Liquid Oxygen (LOX)/Hydrogen, LOX/Methane, and LOX/Kerosene propellants.

Gradl, Paul R.↗
NASA NTRS2009

Sea-Level Flight Demonstration and Altitude Characterization of a LO2/LCH4 Based Accent Propulsion Lander

A 1,500 lbf thrust-class liquid oxygen (LO2)/Liquid Methane (LCH4) rocket engine was developed and tested at both sea-level and simulated altitude conditions. The engine was fabricated by Armadillo Aerospace (AA) in collaboration with NASA Johnson Space Center. Sea level testing was conducted at Armadillo Aerospace facilities at Caddo Mills, TX. Sea-level tests were conducted using both a static horizontal test bed and a vertical take-off and landing (VTOL) test bed capable of lift-off and hover-flight in low atmosphere conditions. The vertical test bed configuration is capable of throttling the engine valves to enable liftoff and hover-flight. Simulated altitude vacuum testing was conducted at NASA Johnson Space Center White Sands Test Facility (WSTF), which is capable of providing altitude simulation greater than 120,000 ft equivalent. The engine tests demonstrated ignition using two different methods, a gas-torch and a pyrotechnic igniter. Both gas torch and pyrotechnic ignition were demonstrated at both sea-level and vacuum conditions. The rocket engine was designed to be configured with three different nozzle configurations, including a dual-bell nozzle geometry. Dual-bell nozzle tests were conducted at WSTF and engine performance data was achieved at both ambient pressure and simulated altitude conditions. Dual-bell nozzle performance data was achieved over a range of altitude conditions from 90,000 ft to 50,000 ft altitude. Thrust and propellant mass flow rates were measured in the tests for specific impulse (Isp) and C* calculations.

Collins, Jacob↗
NASA NTRS2004

High Strength and Compatible Aluminum Alloy for Hydrogen-Peroxide Fuel Tanks

This paper describes the development of a new high strength and Hydrogen Peroxide (HP) propellant compatible aluminum alloy for NASA Hyper-X vehicle's fuel tanks and structures. The tensile strength of the new alloy is more than 3 times stronger than the conventional 5254 alloy while it still maintains HP compatibility similar to 5254 (Class 1 category). The alloy development strategy consists of selecting certain rare earth and transition metals, with unique electrochemical properties, that will not act as catalysts to decompose liquid HP at the atomic level. Such elements will added to the aluminum alloy and the mixture will be cast and rolled into thin sheet metals. Test coupons are machined from sheet metals for HP long-term exposure testing and mechanical properties testing. In addition, the ability to weld the new alloy using Friction Stir Welding has also been explored. Currently, aluminum alloy 5254 is the state-of-the-art material for HP storage, but its yield strength is very low (420 ksi) and may not be suitable for the development of light-weight fuel tanks for Hyper-X vehicles. The new high strength and HP compatible alloy could represent an enabling material technology for NASA's Hyper-X vehicles, where flight weight reduction is a critical requirement. These X-planes are currently under studied as air-breathing hypersonic research vehicles featuring a lifting body configuration with a Rocket Based Combined Cycle (RBCC) engine system.

Lee, Jonathan A.↗
NASA NTRS1954

Preliminary Investigation of Performance and Starting Characteristics of Liquid Fluorine : Liquid Oxygen Mixtures with Jet Fuel

The performance of jet fuel with an oxidant mixture containing 70 percent liquid fluorine and 30 percent liquid oxygen by weight was investigated in a 500-pound-thrust engine operating at a chamber pressure of 300 pounds per square inch absolute. A one-oxidant-on-one-fuel skewed-hole impinging-jet injector was evaluated in a chamber of characteristic length equal to 50 inches. A maximum experimental specific impulse of 268 pound-seconds per pound was obtained at 25 percent fuel, which corresponds to 96 percent of the maximum theoretical specific impulse based on frozen composition expansion. The maximum characteristic velocity obtained was 6050 feet per second at 23 percent fuel, or 94 percent of the theoretical maximum. The average thrust coefficient was 1.38 for the 500-pound thrust combustion-chamber nozzle used, which was 99 percent of the theoretical (frozen) maximum. Mixtures of fluorine and oxygen were found to be self-igniting with jet fuel with fluorine concentrations as low as 4 percent, when low starting propellant flow rated were used.

ENGINES, ROCKET↗
NASA NTRS1999

Non-Toxic Reaction Control System for the Reusable First Stage Vehicle

This paper presents the Boeing Reusable Space Systems vision of a Reaction Control System (RCS) for the Reusable First Stage (RFS) being considered as a replacement for the Solid Rocket Booster for the Space Shuttle. The requirement is to,achieve reliable vehicle control during the upper atmospheric portion of the RFS trajectory while enabling more efficient ground operations, unhindered by constraints caused by operating with highly toxic RCS propellants. Boeing's objective for this effort is to develop a safer, more efficient and environmentally friendly RCS design approach that is suitable for the RFS concept of operations, including a low cost, efficient turnaround cycle. The Boeing RCS concept utilizes ethanol and liquid oxygen in place of the highly toxic, suspected carcinogen, ozone- depleting mono-methyl-hydrazine and highly toxic nitrogen tetroxide. The Space Shuttle Upgrade program, under the leadership of the NASA Johnson Space Flight Center, is currently developing liquid oxygen and ethanol (ethyl alcohol) technology for use as non-toxic orbital maneuvering system (OMS) and RCS. The development of this liquid oxygen and ethanol technology for the Space Shuffle offers a significant leverage to select much of the same technology for the RFS program. There are significant design and development issues involved with bringing this liquid oxygen and ethanol technology to a state of maturity suitable for an operational RCS, The risks associated with a new LOX and Ethanol RCS are mitigated by maintaining kerosene and hydrogen peroxide RCS technology as an alternative. These issues, presented within this paper, include managing the oxygen supply and achieving reliable ignition in the short pulse mode of engine operation. Performance, reliability and operations requirements are presented along with a specific RCS design concept to satisfying these requirements. The work reported in this paper was performed under NASA Marshall Space Flight Center Contract to define Reusable First Stage design concepts for the Space Shuttle.

Keith, E. L.↗
NASA NTRS2007

Infrared Camera Characterization of Bi-Propellant Reaction Control Engines during Auxiliary Propulsion Systems Tests at NASA's White Sands Test Facility in Las Cruces, New Mexico

This paper describes the application of a FUR Systems A40M infrared (IR) digital camera for thermal monitoring of a Liquid Oxygen (LOX) and Ethanol bi-propellant Reaction Control Engine (RCE) during Auxiliary Propulsion System (APS) testing at the National Aeronautics & Space Administration's (NASA) White Sands Test Facility (WSTF) near Las Cruces, New Mexico. Typically, NASA has relied mostly on the use of ThermoCouples (TC) for this type of thermal monitoring due to the variability of constraints required to accurately map rapidly changing temperatures from ambient to glowing hot chamber material. Obtaining accurate real-time temperatures in the JR spectrum is made even more elusive by the changing emissivity of the chamber material as it begins to glow. The parameters evaluated prior to APS testing included: (1) remote operation of the A40M camera using fiber optic Firewire signal sender and receiver units; (2) operation of the camera inside a Pelco explosion proof enclosure with a germanium window; (3) remote analog signal display for real-time monitoring; (4) remote digital data acquisition of the A40M's sensor information using FUR's ThermaCAM Researcher Pro 2.8 software; and (5) overall reliability of the system. An initial characterization report was prepared after the A40M characterization tests at Marshall Space Flight Center (MSFC) to document controlled heat source comparisons to calibrated TCs. Summary IR digital data recorded from WSTF's APS testing is included within this document along with findings, lessons learned, and recommendations for further usage as a monitoring tool for the development of rocket engines.

Holleman, Elizabeth↗
NASA NTRS1999

Non-Toxic Reaction Control System for the Reusable First Stage Vehicle

This paper presents the Boeing Reusable Space Systems vision of a Reaction Control System (RCS) for the Reusable First Stage (RFS) being considered as a replacement for the Solid Rocket Booster for the Space Shuttle. The requirement is to achieve reliable vehicle control during the upper atmospheric portion of the RFS trajectory while enabling more efficient ground operations, unhindered by constraints caused by operating with highly toxic RCS propellants. Boeing's objective for this effort is to develop a safer, more efficient and environmentally friendly RCS design approach that is suitable for the RFS concept of operations, including a low cost, efficient turnaround cycle. The Boeing RCS concept utilizes ethanol and liquid oxygen in place of the highly toxic, suspected carcinogen, ozone-depleting mono-methyl-hydrazine and highly toxic nitrogen tetroxide. The Space Shuttle Upgrade program, under the leadership of the NASA Johnson Space Flight Center, is currently developing liquid oxygen and ethanol (ethyl alcohol) technology for use as non-toxic orbital maneuvering system (OMS) and RCS. The development of this liquid oxygen and ethanol technology for the Space Shuttle offers a significant leverage to select much of the same technology for the RFS program. There are significant design and development issues involved with bringing this liquid oxygen and ethanol technology to a state of maturity suitable for an operational RCS. The risks associated with a new LOX and Ethanol RCS are mitigated by maintaining kerosene and hydrogen peroxide RCS technology as an alternative. These issues, presented within this paper, include managing the oxygen supply and achieving reliable ignition in the short pulse mode of engine operation. Performance, reliability and operations requirements are presented along with a specific RCS design concept to satisfying these requirements. The work reported in this paper was performed under NASA Marshall Space Flight Center Contract Number NAS8-97272 to define Reusable First Stage design concepts for the Space Shuttle.

Keith, E. L.↗
NASA NTRSDate not supplied

NASA’s Rotating Detonation Rocket Engine Development

The Rotating Detonation Rocket Engine has maintained steady development at NASA with many staggering performance advantages demonstrated to date over the state-of-the-art (SOA). The implementation of additive manufacturing and specialized NASA developed alloys have enabled rapid maturation of the technology. Several hot fire test projects have been successfully conducted at Marshall Space Flight Center under an early career initiative project funded by NASA Space Technology Mission Directorate. In addition, a new start Technology Demonstration Mission (TDM) project has been funded to investigate challenges relating to integration of turbomachinery with an RDRE thrust chamber assembly. This engine system demonstration will leverage a methane/oxygen single shaft turbopump with fuel rich gas generator and a 10,000 lbf thrust chamber assembly. The configuration was down selected based on feedback from both US industry collaborators and power balance trades in combination with technical feasibility. To date, industry has identified several use cases for RDRE ranging from thruster to primary launch vehicle propulsion. A wide range of fuel and oxidizers were also identified including but not limited to Methane, Kerosene and other liquid hydrocarbon (LH) fuels, and hydrogen. Recent work at NASA and in partnership with NASA has investigated these major fuels of interest with oxygen, air, and hydrogen peroxide (HTP) for various applications. NASA Marshall has already investigated the use of hydrogen/oxygen, methane/oxygen, kerosene/oxygen, and has plans in partnership with industry and academia to investigate LH/air and LH/HTP. In addition to propellants, hardware geometry has been investigated with some critical lessons learned toward greater theoretical performance over the SOA. To this end, several experimental and computational activities are ongoing to further advance the RDRE towards flight missions. Given the rate of advancement, it is highly likely the technology will be flown in space mission in the coming decade. This work documents and overviews many of these investigations and overviews NASA’s future plans for the technology maturation.

Thomas Teasley↗