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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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The capabilities and characteristics of the RL10 rocket engine are examined. The engine model history is presented. The RL10 derivatives are also outlined. The presentation is represented by viewgraphs.
In July of 1999 two linear aerospike rocket engines will power the first flight of NASA's X-33 advanced technology demonstrator. A successful X-33 flight test program will validate the aerospike nozzle concept, a key technical feature of Lockheed Martin's VentureStar(trademark) reusable launch vehicle. The aerospike received serious consideration for NASA's current space shuttle, but was eventually rejected in 1969 in favor of high chamber pressure bell engines, in part because of perceived technical risk. The aerospike engine (discussed below) has several performance advantages over conventional bell engines. However, these performance advantages are difficult to validate by ground test. The space shuttle, a multibillion dollar program intended to provide all of NASA's future space lift could not afford the gamble of choosing a potentially superior though unproven aerospike engine over a conventional bell engine. The X-33 demonstrator provides an opportunity to prove the aerospike's performance advantage in flight before commiting to an operational vehicle.
A LOX/Methane rocket engine was developed for a prototype terrestrial lander and then used to fly the lander at Johnson Space Center. The development path of this engine is outlined, including unique items such as variable acoustic damping and variable film cooling.
Algol rocket engine and motor ground support equipment for Little Joe II launch vehicle
Temperature and density measurements in base region flow field of clustered rocket engine model using electron beam technique
Rocket engine record, x-15 aircraft, speed and altitude of research aircraft
Absorption, scattering, and extinction parameters of small spherical metal particles used as seeding agents to control radiant heat transfer in gaseous nuclear rocket engines
Cooling tests of M-1 rocket engine turbopump with liquid hydrogen
Static and dynamic combustion phenomena effects on grain structural design for solid propellant rocket engine
Third 260-inch diameter solid propellant rocket engine test firing
Much deserved attention is given to the feats of innovation that allow humans to live in space and robotic explorers to beam never-before-seen images back to Earth. In the background of these accomplishments is a technology that makes it all possible the rockets that propel NASA s space exploration efforts skyward. Marshall Space Flight Center has been at the heart of the Agency s rocketry and spacecraft propulsion efforts since its founding in 1960. Located at the Redstone Arsenal near Huntsville, Alabama, the Center has a legacy of success stretching back to the Saturn rockets that carried the Apollo astronauts into space. Even before Marshall was established, Redstone was the site of significant advances in American rocketry under the guidance of famous rocket engineer Werner Von Braun; these included the Juno I rocket that successfully carried the United States first satellite, Explorer 1, into orbit in 1958. And from the first orbital test flight of the Space Shuttle Columbia through the final flights of the shuttle program this year, these vehicles have been enabled by the solid rocket boosters, external tank, and orbiter main engines created at Marshall. Today, Marshall continues to host innovation in rocket and spacecraft propulsion at state-of-the-art facilities such as the Propulsion Research Laboratory. Like many of its past successes, some of the Center s current advancements are being made with the help of private industry partners. The efforts have led not only to new propulsion technologies, but to terrestrial benefits in a seemingly unrelated field in this case, firefighting.
Lubrication characteristics of bearing steel in liquid oxygen in rocket engines
Analysis of heat transfer in rocket engine thrust chambers that use convergent-divergent nozzles
The Nuclear Thermal Rocket (NTR) represents the next evolutionary step in cryogenic liquid rocket engines. Deriving its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core, the NTR can generate high thrust at a specific impulse of approx. 900 seconds or more - twice that of today's best chemical rockets. In FY'11, as part of the AISP project, NASA proposed a Nuclear Thermal Propulsion (NTP) effort that envisioned two key activities - "Foundational Technology Development" followed by system-level "Technology Demonstrations". Five near-term NTP activities identified for Foundational Technology Development became the basis for the NCPS project started in FY'12 and funded by NASA's AES program. During Phase 1 (FY'12-14), the NCPS project was focused on (1) Recapturing fuel processing techniques and fabricating partial length "heritage" fuel elements for the two candidate fuel forms identified by NASA and the DOE - NERVA graphite "composite" and the uranium dioxide (UO2) in tungsten "cermet". The Phase 1 effort also included: (2) Engine Conceptual Design; (3) Mission Analysis and Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable and Sustainable NTP Development Strategy. During FY'14, a preliminary plan for DDT&E was outlined by GRC, the DOE and industry for NASA HQ that involved significant system-level demonstration projects that included GTD tests at the NNSS, followed by a FTD mission. To reduce development costs, the GTD and FTD tests use a small, low thrust (approx. 7.5 or 16.5 klbf) engine. Both engines use graphite composite fuel and a "common" fuel element design that is scalable to higher thrust (approx. 25 klbf) engines by increasing the number of elements in a larger diameter core that can produce greater thermal power output. To keep the FTD mission cost down, a simple "1-burn" lunar flyby mission was considered along with maximizing the use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10-B2 engine and Delta Cryogenic Second Stage) to further ensure affordability. This paper provides a preliminary NASA, DOE and industry assessment of what is required - the key DDT&E activities, development options, and the associated schedule - to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.
Liquan program designed to analyze liquid propellant rocket engine static firing data - avoiding blackbox computer programs
Surveyor Vernier thrust chamber assembly - bipropellant rocket engine, flow control, control interfaces, and test firing
Ultrasonic pulse technique for measuring average gas temperature in nuclear rocket engine - sound propagation and environmental studies
Solid propellant rocket engine performance computer programs using group transformation method