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At least 1,189 records · Page 66
All sourcesNASA NTRSDOE OSTI
NASA NTRS2010

Planning for Plume Diagnostics for Ground Testing of J-2X Engines at the SSC

John C. Stennis Space Center (SSC) is the premier test facility for liquid rocket engine development and certification for the National Aeronautics and Space Administration (NASA). Therefore, it is no surprise that the SSC will play the most prominent role in the engine development testing and certification for the J-2X engine. The Pratt & Whitney Rocketdyne J-2X engine has been selected by the Constellation Program to power the Ares I Upper Stage Element and the Ares V Earth Departure Stage in NASA s strategy of risk mitigation for hardware development by building on the Apollo program and other lessons learned to deliver a human-rated engine that is on an aggressive development schedule, with first demonstration flight in 2010 and human test flights in 2012. Accordingly, J-2X engine design, development, test, and evaluation is to build upon heritage hardware and apply valuable experience gained from past development and testing efforts. In order to leverage SSC s successful and innovative expertise in the plume diagnostics for the space shuttle main engine (SSME) health monitoring,1-10 this paper will present a blueprint for plume diagnostics for various proposed ground testing activities for J-2X at SSC. Complete description of the SSC s test facilities, supporting infrastructure, and test facilities is available in Ref. 11. The A-1 Test Stand is currently being prepared for testing the J-2X engine at sea level conditions. The A-2 Test Stand is currently being used for testing the SSME and may also be used for testing the J-2X engine at sea level conditions in the future. Very recently, ground-breaking ceremony for the new A-3 rocket engine test stand took place at SSC on August 23, 2007. A-3 is the first large - scale test stand to be built at the SSC since the A and B stands were constructed in the 1960s. The A-3 Test Stand will be used for testing J-2X engines under vacuum conditions simulating high altitude operation at approximately 30,480 m (100,000 ft). To achieve the simulated altitude environment, chemical steam generators using isopropyl alcohol, LOX, and RELEASED - Printed documents may be obsolete; validate prior to use. water would run for the duration of the test and would generate approximately 2096 Kg/s of steam to reduce pressure in the test cell and downstream of the engine. The testing at the A-3 Test Stand is projected to begin in late 2010, meanwhile the J-2X component testing on A-1 is scheduled to begin later this year.

SaintCyr, William W.↗
NASA NTRS1966

Some Observations on Exhaust Recirculation from Clustered Rocket Nozzles

A short-duration rocket testing technique which exactly duplicates the combustion flows of multi-engine rocket boosters at high altitudes is described. The application of this technique to the investigation of the reverse flow characteristics of a 4-engine rocket cluster model is discussed. The use of a comparatively new diagnostic tool, the electron beam density probe, in mapping the recirculating flow field of a rocket cluster is described and experimental density results presented. Finally, the use of a reflection plane for the examination of reverse flow phenomena, whereby a plane of symmetry is replaced by a solid surface on which measurements may be made,is evaluated and experimental data presented.

Hendershot, K. C.↗
NASA NTRS1973

Flight experience with Apollo spacecraft propulsion systems

Apollo 17 ended the most successful application of rocket propulsion systems in man's history. A total of 23 developmental and manned operational flights were made. Seven hundred and sixty-three spacecraft rocket engines were flown in the program. Over 6 h of manned rocket flights were logged by the spacecraft propulsion systems and approximately one million rocket engine firings were made. One engine failure was encountered on an early unmanned flight as a result of a failure in the guidance programmer which caused the engine to operate in a manner known to cause failures. Numerous operational problems and malfunctions were observed; however, system and component redundancy prevented loss of mission objectives and never jeopardized crew safety. Performance of all systems was usually nominal and most problems were merely nuisances. This paper will present some highlights of Apollo propulsion performance and will provide a bibliography of all flight results.

Thibodaux, J. G., Jr.↗
NASA NTRS1987

Low Thrust Space Propulsion

An overview of low rocket engine propulsion concepts for space missions is presented. Chemical and electrical rocket engines are shown. Animation illustrates propulsion applications.

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NASA NTRS1994

Real-time sensor data validation

This report describes the status of an on-going effort to develop software capable of detecting sensor failures on rocket engines in real time. This software could be used in a rocket engine controller to prevent the erroneous shutdown of an engine due to sensor failures which would otherwise be interpreted as engine failures by the control software. The approach taken combines analytical redundancy with Bayesian belief networks to provide a solution which has well defined real-time characteristics and well-defined error rates. Analytical redundancy is a technique in which a sensor's value is predicted by using values from other sensors and known or empirically derived mathematical relations. A set of sensors and a set of relations among them form a network of cross-checks which can be used to periodically validate all of the sensors in the network. Bayesian belief networks provide a method of determining if each of the sensors in the network is valid, given the results of the cross-checks. This approach has been successfully demonstrated on the Technology Test Bed Engine at the NASA Marshall Space Flight Center. Current efforts are focused on extending the system to provide a validation capability for 100 sensors on the Space Shuttle Main Engine.

Bickmore, Timothy W.↗
NASA NTRS2003

Development of Efficient Real-Fluid Model in Simulating Liquid Rocket Injector Flows

The characteristics of propellant mixing near the injector have a profound effect on the liquid rocket engine performance. However, the flow features near the injector of liquid rocket engines are extremely complicated, for example supercritical-pressure spray, turbulent mixing, and chemical reactions are present. Previously, a homogeneous spray approach with a real-fluid property model was developed to account for the compressibility and evaporation effects such that thermodynamics properties of a mixture at a wide range of pressures and temperatures can be properly calculated, including liquid-phase, gas- phase, two-phase, and dense fluid regions. The developed homogeneous spray model demonstrated a good success in simulating uni- element shear coaxial injector spray combustion flows. However, the real-fluid model suffered a computational deficiency when applied to a pressure-based computational fluid dynamics (CFD) code. The deficiency is caused by the pressure and enthalpy being the independent variables in the solution procedure of a pressure-based code, whereas the real-fluid model utilizes density and temperature as independent variables. The objective of the present research work is to improve the computational efficiency of the real-fluid property model in computing thermal properties. The proposed approach is called an efficient real-fluid model, and the improvement of computational efficiency is achieved by using a combination of a liquid species and a gaseous species to represent a real-fluid species.

Cheng, Gary↗
NASA NTRS1992

Reliability evaluation methodology for NASA applications

Liquid rocket engine technology has been characterized by the development of complex systems containing large number of subsystems, components, and parts. The trend to even larger and more complex system is continuing. The liquid rocket engineers have been focusing mainly on performance driven designs to increase payload delivery of a launch vehicle for a given mission. In otherwords, although the failure of a single inexpensive part or component may cause the failure of the system, reliability in general has not been considered as one of the system parameters like cost or performance. Up till now, quantification of reliability has not been a consideration during system design and development in the liquid rocket industry. Engineers and managers have long been aware of the fact that the reliability of the system increases during development, but no serious attempts have been made to quantify reliability. As a result, a method to quantify reliability during design and development is needed. This includes application of probabilistic models which utilize both engineering analysis and test data. Classical methods require the use of operating data for reliability demonstration. In contrast, the method described in this paper is based on similarity, analysis, and testing combined with Bayesian statistical analysis.

Taneja, Vidya S.↗
NASA NTRS1997

Demonstration of Oxygen and Carbon Monoxide Propellants for Mars Missions

Currently, proposed planetary exploration missions must be small, with low costs and a short development time. Relatively high-risk technologies are being accepted for such missions if they meet these guidelines. For a Mars sample-return mission, one of the higher risk technologies is the use of return propellants produced from indigenous materials such as the Martian atmosphere. This consists of 96 percent carbon dioxide, which can be processed into oxygen and carbon monoxide. This year, the NASA Lewis Research Center completed the experimental evaluation and subscale technology development of an oxygen/carbon monoxide propellant combination. Previous research included ignition characterization, combustion performance, and heat transfer characterization with gaseous propellants at room temperature. In this year s tests, we studied the ignition characteristics and combustion of oxygen and carbon monoxide at near liquid temperatures. The mixture ratio boundaries for oxygen and carbon monoxide were determined as a function of propellant temperature in a spark torch igniter. With both propellants at room temperature, the ignition range was between 0.50 and 1.44; and with both propellants chilled to near-liquid temperatures, it was between 2.4 and 3.1. Statistical analysis of the mean value of the ignition boundaries provided models that describe the combination of oxygen temperature, carbon monoxide temperature, and mixture ratio that resulted in ignition. This range is the larger boxed area shown in the figure. The smaller boxed area indicates the range at which there is a 90-percent confidence that ignition will occur. The relatively small range at only 90-percent confidence indicates that using the oxygen/carbon monoxide combination as its own ignition source may not be the best design for a remote engine operating on Mars. Tests also were performed in a simulated small rocket engine that used oxygen/hydrogen combustion gases as the ignition source for oxygen/carbon monoxide. In these experiments, the oxygen/carbon monoxide was successfully ignited in eight of eight tests at a mixture ratio of 0.52. In addition, the oxygen/carbon monoxide maintained steady combustion after the oxygen/hydrogen ignition source was removed, verifying that the oxygen/carbon monoxide rocket engine should continue to be included in mission plans as return propulsion from Mars.

Linne, Diane L.↗
NASA NTRS1988

Ceramic matrix composites in simulated SSME environments

Future Space Shuttle Main Engine (SSME)-type rocket engines can benefit from the use of fiber-reinforced ceramix matrix composites (FRCMC). Ceramics reinforced with long continuous fibers exhibit improved tolerance to severe thermomechanical and environmental exposures. An in-house NASA-Lewis program to evaluate the durability of FRCMC in simulated SSME environments is described. Primary tests involve multiple (one second) exposures of FRCMC specimens in a hydrogen/oxygen rocket test rig. This rig generates surface heating rates of 1000 to 2500 C/second. The FRCMC durability evaluation involves measurement of retained strength as a function of thermal shock severity and number of upshock cycles. Preliminary test results for monolithic silicon nitride (Si3N4) and silicon carbide (SiC), and one type of silicon based FRCMC, are presented. The test data are examined in terms of simple thermal shock theory.

Herbell, Thomas P.↗
NASA NTRS1969

Two-axis winch installer for heavy ducts in confined space

Two-axis winching and traversing device is used for installing liquid-propellant rocket-engine fuel and liquid oxygen suction ducts between the valves and the rocket engine on a test stand. The device raises and maneuvers the duct into the required position where it can be safely installed by mechanics.

Cox, E. F.↗
NASA NTRS1972

Forebody and vertical stabilizer effects on directional stability of a reusable LOX/RP (061) booster AR 12161-2

Results are presented of a wind tunnel test on the directional stability of space shuttle booster configurations. The test was conducted at the 14-inch trisonic tunnel starting 6 December 1971 and continued through 11 December 1971 for a total of 66 occupancy hours. Configurations tested included a cylindrical body with two axisymmetrical noses, one with and without canopy, one delta wing, located in two positions, five vertical tails (including a V tail), two having split rudders, ventral fins, two sets of chines, three airbreathing engine pods, and rocket engine shrouds. The model scale was 0.003366.

Roth, C. E.↗
NASA NTRS1977

Liquid Rocket Lines, Bellows, Flexible Hoses, and Filters

Fluid-flow components in a liquid propellant rocket engine and the rocket vehicle which it propels are interconnected by lines, bellows, and flexible hoses. Elements involved in the successful design of these components are identified and current technologies pertaining to these elements are reviewed, assessed, and summarized to provide a technology base for a checklist of rules to be followed by project managers in guiding a design or assessing its adequacy. Recommended procedures for satisfying each of the design criteria are included.

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NASA NTRS1978

The Space Shuttle

The Space Transportation System (STS) consisting of the Space Shuttle, Spacelab and the Upper Stages is described, together with the Shuttle's projected missions, test programs and costs. The Orbiter, due to be launched late in 1979, is roughly the size of a DC-9, with weight dry empty of about 75,000 kg, and is reusable, as are the two external fuel tanks. The shape and size of the cargo bay (unobstructed cylindrical compartment, 4.6 m in diameter and 18.3 m long), the engine structure (three rocket engines, fueled by liquid oxygen and liquid hydrogen, each capable of developing 1,668,080 N of thrust), the atmospheric condition in the Orbiter's cabin at sea level pressure, the surface insulation material of coated silica-fiber tiles and reinforced carbon-carbon material with a protection capability of up to 1650 C are discussed in detail. The Spacelab, designed by ESA to fit in the Orbiter's cargo bay for scientific research, and the Upper Stages, intended for inserting payloads in high-energy earth orbits, are also analyzed, as are the two-frequency band communication systems and the on board hardware.

Malkin, M. S.↗
NASA NTRS1991

Lightweight ceramic components for propulsion applications

It is demonstrated that two ceramic engine components of reduced weight can be efficiently produced and employed at higher combustion gas temperatures. The engine components, containing internal cooling and hydraulic passages, are fabricated in a process involving platelet technology, tape casting, and multilayer ceramic packaging techniques. Traditional metal platelet design techniques are used to fabricate components from ceramic tape which are then laminated and processed to densities of close to 100 percent. A methane cooled regenerator for an air turbo ramjet engine and a rocket engine bipropellant injector are fabricated from silicon nitride. The two components are tested under conditions similar to the pressurization, thrusts, and temperatures encountered in a combustion gas environment. The survivability and leak-free nature of the components demonstrate that more complex ceramic engine components can be fabricated using these techniques.

Franklin, J. E.↗
NASA NTRS1998

Key Issues for Aerospace Applications of Ceramic Matrix Composites

Ceramic matrix composites (CMC) offer significant advantages for future aerospace applications including turbine engine and liquid rocket engine components, thermal protection systems, and "hot structures". Key characteristics which establish ceramic matrix composites as attractive and often enabling choices are strength retention at high temperatures and reduced weight relative to currently used metallics. However, due to the immaturity of this class of materials which is further compounded by the lack of experience with CMC's in the aerospace industry, there are significant challenges involved in the development and implementation of ceramic matrix composites into aerospace systems. Some of the more critical challenges are attachment and load transfer methodologies; manufacturing techniques, particularly scale up to large and thick section components; operational environment resistance; damage tolerance; durability; repair techniques; reproducibility; database availability; and the lack of validated design and analysis tools. The presentation will examine the technical issues confronting the application of ceramic matrix composites to aerospace systems and identify the key material systems having potential for substantial payoff relative to the primary requirements of light weight and reduced cost for future systems. Current programs and future research opportunities will be described in the presentation which will focus on materials and processes issues.

Clinton, R. G., Jr.↗
NASA NTRS2001

National Rocket Propulsion Materials Plan: A NASA, Department of Defense, and Industry Partnership

NASA, Department of Defense, and rocket propulsion industry representatives are working together to create a national rocket propulsion materials development roadmap. This "living document" will facilitate collaboration among the partners, leveraging of resources, and will be a highly effective tool for technology development planning. The structuring of the roadmap, and development plan, which will combine the significant efforts of the Integrated High Payoff Rocket Propulsion Technology (IHPRPT) Program, and NASA's Integrated Space Transportation Plan (ISTP), is being lead by the IHPRPT Materials Working Group (IMWG). The IHPRPT Program is a joint DoD, NASA, and industry effort to dramatically improve the nation's rocket propulsion capabilities. This phased program is structured with increasingly challenging goals focused on performance, reliability, and cost to effectively double rocket propulsion capabilities by 2010. The IHPRPT program is focused on three propulsion application areas: Boost and Orbit Transfer (both liquid rocket engines and solid rocket motors), Tactical, and Spacecraft. Critical to the success of this initiative is the development and application of advanced materials, processes, and manufacturing technologies. NASA's ISTP is a comprehensive strategy focusing on the aggressive safety, reliability, and affordability goals for future space transportation systems established by the agency. Key elements of this plan are the 2 nd and 3 d Generation Reusable Launch Vehicles (RLV). The affordability and safety goals of these generational systems are, respectively, 10X cheaper and 100X safer by 2010, and 100X cheaper and 10,000X safer by 2025. Accomplishment of these goals requires dramatic and sustained breakthroughs, particularly in the development and the application of advanced material systems. The presentation will provide an overview of the IHPRPT materials initiatives, NASA's 2nd and 3 rd Generation RLV propulsion materials projects, and the approach for the development of the national rocket propulsion materials roadmap.

Clinton, Raymond G., Jr.↗
NASA NTRS2003

NASA Tech Briefs, June 2003

Topics covered include: Nulling Infrared Radiometer for Measuring Temperature; The Ames Power Monitoring System; Hot Films on Ceramic Substrates for Measuring Skin Friction; Probe Without Moving Parts Measures Flow Angle; Detecting Conductive Liquid Leaking from Nonconductive Pipe; Adaptive Suppression of Noise in Voice Communications; High-Performance Solid-State W-Band Power Amplifiers; Microbatteries for Combinatorial Studies of Conventional Lithium-Ion Batteries; Correcting for Beam Aberrations in a Beam-Waveguide Antenna; Advanced Rainbow Solar Photovoltaic Arrays; Metal Side Reflectors for Trapping Light in QWIPs; Software for Collaborative Engineering of Launch Rockets; Software Assists in Extensive Environmental Auditing; Software Supports Distributed Operations via the Internet; Software Estimates Costs of Testing Rocket Engines; yourSky: Custom Sky-Image Mosaics via the Internet; Software for Managing Inventory of Flight Hardware; Lower-Conductivity Thermal-Barrier Coatings; Process for Smoothing an Si Substrate after Etching of SiO2; Flexible Composite-Material Pressure Vessel; Treatment to Destroy Chlorohydrocarbon Liquids in the Ground; Noncircular Cross Sections Could Enhance Mixing in Sprays; Small, Untethered, Mobile Roots for Inspecting Gas Pipes; Paint-Overspray Catcher; Preparation of Regular Specimens for Atom Probes; Inverse Tomo-Lithography for Making Microscopic 3D Parts; Predicting and Preventing Incipient Flameout in Combustors; MEMS-Based Piezoelectric/Electrostatic Inchworm Actuator; Metallized Capillaries as Probes for Raman Spectroscopy; Adaptation of Mesoscale Weather Models to Local Forecasting; Aerodynamic Design using Neural Networks; Combining Multiple Gyroscope Outputs for Increased Accuracy; and Improved Collision-Detection Method for Robotic Manipulator.

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