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Jet engine being set up for testing with technicians from Propulsion System Laboratory (PSL). Image modified to include photographer taking the photo.

Jet engine being set up for testing with technicians from Propulsion System Laboratory (PSL). Image modified to include photographer taking the photo.

A cooled-gas pyrometer designed for application in a hypersonic research engine program was fabricated and tested. Design and operational considerations and calibration data are presented. The probe was tested in a rocket-engine exhaust stream operating at Mach 2 and 2300 K. Test temperature measurements agreed to within 2 percent with a radiation shielded thermocouple probe.

One of the most important properties of aviation fuels for spark-ignition engines is their knock rating. The CFR engine tests of fuels of 87 octane and above does not always correspond entirely to the actual behavior of these fuels in the airplane engine. A method is therefore developed which, in contrast to the octane number determination, permits a testing of the fuel under various temperatures and fuel mixture conditions. The following reference fuels were employed: 1) Primary fuels; isooctane and n-heptane; 2) Secondary fuels; pure benzene and synthetic benzine.

The Augmented Deflector Exhaust Nozzle (ADEN) exhaust system was tested in the PSL-3 altitude chamber at the NASA Lewis Research Center in order to evaluate aerodynamic performance, cooling-system effectiveness, and mechanical operation at flight-type conditions. The ADEN, a flight-weight, two-dimensional, thrust-vectoring nozzle, was successfully tested on the F404 engine using a remote engine control system for automatic or manual setting of the throat-area control and available fan air for the nozzle internal cooling system. Throughout the test, the ADEN performed with no adverse effects on the engine or augmentor operation.

The performance of a turbocharged rotary engine at power levels above 75 kW (100 hp) was studied. A twin rotor turbocharged Mazda engine was tested at speeds of 3000 to 6000 rpm and boost pressures to 7 psi. The NASA developed combustion diagnostic instrumentation was used to quantify indicated and pumping mean effect pressures, peak pressure, and face to face variability on a cycle by cycle basis. Results of this testing showed that a 5900 rpm a 36 percent increase in power was obtained by operating the engine in the turbocharged configuration. When operating with lean carburetor jets at 105 hp (78.3 kW) and 4000 rpm, a brake specific fuel consumption of 0.45 lbm/lb-hr was measured.

No Abstract Available

No Abstract Available

Ultra-Efficient Engine Technology (UEET), Proof of Concept Compressor, Advanced Compressor Casing Treatment testing; bearing housing and shaft during build-up

Ultra-Efficient Engine Technology (UEET) Proof of Concept Compressor, Advanced Compressor Casing Treatment testing, First Research Configuration (concentric grooves)

Jet engine being set up for testing

Jet engine being set up for testing

Ultra-Efficient Engine Technology (UEET), Proof of Concept Compressor, Advanced Compressor Casing Treatment testing; bearing housing and shaft during build-up

This report details the use of infrared imaging for temperature measurement and thermal pattern determination during simulated altitude engine testing in the NASA Lewis Propulsion Systems Laboratory. Three identical argon-cooled imaging systems were installed in the facility exhaust collector behind sapphire windows to look at engine internal surfaces. The report describes the components of each system, presents the specifics of the complicated installation, and explains the operation of the systems during engine testing. During the program, several problems emerged, such as argon contamination system, component overheating, cracked sapphire windows, and other unexplained effects. This report includes a summary of the difficulties as well as the solutions developed. The systems performed well, considering they were in an unusually harsh exhaust environment. Both video and digital data were recorded, and the information provided valuable material for the engineers and designers to quickly make any necessary design changes to the engine hardware cooling system. The knowledge and experience gained during this program greatly simplified the installation and use of the systems during later test programs in the facility. The infrared imaging systems have significantly enhanced the measurement capabilities of the facility, and have become an outstanding and versatile testing resource in the Propulsion Systems Laboratory.

Five firings of the Rocketdyne J-2 rocket engine were conducted in Test Cell J-4 of the Large Rocket Facility. The firings were accomplished during test period J4-1801-23 at pressure altitudes ranging from 105,500 to 113,000 ft at engine start to evaluate the effects of predicted Saturn S-II stage pre-fire conditions on J-2 engine start transients. Engine components were thermally conditioned to simulate the expected S-II interstage/engine thermal environment. Satisfactory engine operation was obtained. Accumulated firing duration was 56.3 sec.

A test was conducted by the Air Force Aero Propulsion Laboratory to study the performance of jet engine compressors during engine speed transients. Effective analysis of the test results required accurate, high speed data acquisition during the transients. On-line reduction and display of the performance data was desirable to improve the efficiency of the test operations. This paper describes the software that was designed and implemented to acquire and display the performance data. The software controls data acquisition and calibration equipment, provides on-line calibration of transducers, records the data, reduces and displays performance data (including graphics), and allows the test operator to control these functions. (Author)

This photograph depicts a test firing of an F-1 engine at the F-1 engine test stand in the west test area of the Marshall Space Flight Center. This engine produced 1,500,000 pounds of thrust using liquid oxygen and RP-1, which is a derivative of kerosene. The F-1 engine test stand was constructed in 1963 to assist in the development of the F-1 engine.

Optical fire detector systems for aircraft engine compartments provided by fire vendors were evaluated in the Aircraft Engine Compartment Fire Test Simulator (AENFTS). Their ability to sense JP-4 fires of two sizes and provide fire warnings in a variety of simulated flight environments was monitored along with their freedom from false fire warnings. One systems was identified which provided fire warnings for every test fire and was completely free of false alarms.

INCONEL 718 panels after impact testing in the Ballistics Impact Lab with Titanium Simulated Jet Engine Blade

Many dynamic systems operate in select operating regions, each exhibiting characteristic modes of behavior. It is traditional to employ standard adjustable gain proportional-integral-derivative (PID) loops in such systems where no apriori model information is available. However, for controlling inlet pressure for rocket engine testing, problems in fine tuning, disturbance accommodation, and control gains for new profile operating regions (for research and development) are typically encountered. Because of the capability of capturing I/O peculiarities, using NETS, a back propagation trained neural network is specified. For select operating regions, the neural network controller is simulated to be as robust as the PID controller. For a comparative analysis, the higher order moment neural array (HOMNA) method is used to specify a second neural controller by extracting critical exemplars from the I/O data set. Furthermore, using the critical exemplars from the HOMNA method, a third neural controller is developed using NETS back propagation algorithm. All controllers are benchmarked against each other.

Many dynamic systems operate in select operating regions, each exhibiting characteristic modes of behavior. It is traditional to employ standard adjustable gain proportional-integral-derivative (PID) loops in such systems where no apriori model information is available. However, for controlling inlet pressure for rocket engine testing, problems in fine tuning, disturbance accommodation, and control gains for new profile operating regions (for research and development) are typically encountered. Because of the capability of capturing I/O peculiarities, using NETS, a back propagation trained neural network is specified. For select operating regions, the neural network controller is simulated to be as robust as the PID controller. For a comparative analysis, the higher order moment neural array (HOMNA) method is used to specify a second neural controller by extracting critical exemplars from the I/O data set. Furthermore, using the critical exemplars from the HOMNA method, a third neural controller is developed using NETS back propagation algorithm. All controllers are benchmarked against each other.

Metallic contaminant was found in the liquid oxygen (LOX) pre-valve screen of the shuttle main engine propulsion system on two orbiter vehicles. To investigate the potential for an ignition, NASA Johnson Space Center White Sands Test Facility performed (modified) rotating friction ignition testing in LOX. This testing simulated a contaminant particle in the low-pressure oxygen turbo pump (LPOTP) and the high-pressure oxygen turbo pump (HPOTP) of the shuttle main propulsion system. Monel(R) K-500 and Inconel(R) 718 samples represented the LPOTP and HPOTP materials. Aluminum foil tape and titanium foil represented the contaminant particles. In both the Monel(R) and Inconel(R) material configurations, the aluminum foil tape samples did not ignite after 30 s of rubbing. In contrast, all of the titanium foil samples ignited regardless of the rubbing duration or material configuration. However, the titanium foil ignitions did not propagate to the Monel and Inconel materials.

This photograph depicts a dramatic view of the first test firing of all five F-1 engines for the Saturn V S-IC stage at the Marshall Space Flight Center. The testing lasted a full duration of 6.5 seconds. It also marked the first test performed in the new S-IC static test stand and the first test using the new control blockhouse. The S-IC stage is the first stage, or booster, of a 364-foot long rocket that ultimately took astronauts to the Moon. Operating at maximum power, all five of the engines produced 7,500,000 pounds of thrust. Required to hold down the brute force of a 7,500,000-pound thrust, the S-IC static test stand was designed and constructed with the strength of hundreds of tons of steel and cement, planted down to bedrock 40 feet below ground level. The structure was topped by a crane with a 135-foot boom. With the boom in the up position, the stand was given an overall height of 405 feet, placing it among the highest structures in Alabama at the time. When the Saturn V S-IC first stage was placed upright in the stand , the five F-1 engine nozzles pointed downward on a 1,900 ton, water-cooled deflector. To prevent melting damage, water was sprayed through small holes in the deflector at the rate 320,000 gallons per minute.

This paper discusses simulation technology used to support the testing of rocket propulsion systems by performing high fidelity analyses of feed system components. A generalized multi-element framework has been used to perform simulations of control valve systems. This framework provides the flexibility to resolve the structural and functional complexities typically associated with valve-based high pressure feed systems that are difficult to deal with using traditional Computational Fluid Dynamics (CFD) methods. In order to validate this framework for control valve systems, results are presented for simulations of a cryogenic control valve at various plug settings and compared to both experimental data and simulation results obtained at NASA Stennis Space Center. A detailed unsteady analysis has also been performed for a pressure regulator type control valve used to support rocket engine and component testing at Stennis Space Center. The transient simulation captures the onset of a modal instability that has been observed in the operation of the valve. A discussion of the flow physics responsible for the instability and a prediction of the dominant modes associated with the fluctuations is presented.

The research and development activities reported in this publication were carried out under NASA Aeronautics Research Institute (NARI) funded project entitled "A Fully Nonmetallic Gas Turbine Engine Enabled by Additive Manufacturing." The objective of the project was to conduct evaluation of emerging materials and manufacturing technologies that will enable fully nonmetallic gas turbine engines. The results of the activities are described in three part report. The first part of the report contains the data and analysis of engine system trade studies, which were carried out to estimate reduction in engine emissions and fuel burn enabled due to advanced materials and manufacturing processes. A number of key engine components were identified in which advanced materials and additive manufacturing processes would provide the most significant benefits to engine operation. The technical scope of activities included an assessment of the feasibility of using additive manufacturing technologies to fabricate gas turbine engine components from polymer and ceramic matrix composites, which were accomplished by fabricating prototype engine components and testing them in simulated engine operating conditions. The manufacturing process parameters were developed and optimized for polymer and ceramic composites (described in detail in the second and third part of the report). A number of prototype components (inlet guide vane (IGV), acoustic liners, engine access door) were additively manufactured using high temperature polymer materials. Ceramic matrix composite components included turbine nozzle components. In addition, IGVs and acoustic liners were tested in simulated engine conditions in test rigs. The test results are reported and discussed in detail.

Post-test investigation of a T-700 engine brush seal found regions void of bristles ('yanked out'), regions of bent-over bristles near the inlet, some 'snapped' bristles near the fence, and a more uniform 'smeared' bristle interface between the first and last axial rows of bristles. Several bristles were cut from the brush seal, wax mounted, polished, and analyzed. Metallographic analysis of the bristles near the rub tip showed tungsten-rich phases uniformly distributed throughout the bristle with no apparent change within 1 to 2 micron of the interface except for possibly a small amount of titanium, which would represent a transfer from the rotor. Analysis of the bristle wear face showed nonuniform tungsten, which is indicative of material resolidification. The cut end contained oxides and internal fractures; the worn end was covered with oxide scale. Material losses due to wear and elastoplastic deformation within the shear zone and third-body lubrication effects in the contact zone are discussed.

Testing of a variable-stroke Stirling engine at NASA Lewis has been completed. In support of the DOE Stirling Engine Highway Vehicle Systems Program, the engine was tested for about 70 hours total with both He and H2 as working fluids over a range of pressures and strokes. A direct comparison was made of part-load efficiencies obtained with variable-stroke (VS) and variable-pressure operation. Two failures with the variable-angle swash-plate drive system limited testing to low power levels. These failures are not thought to be caused by problems inherent with the VS concept but do emphasize the need for careful design in the area of the crossheads.

As part of the NASA funded portion of the SP-100 Advanced Technology Program the Space Power Research Engine (SPRE 1) was designed and built to serve as a research tool for evaluation and development of advanced Stirling engine concepts. The SPRE 1 is designed to produce 12.5 kW electrical power when operated with helium at 15 MPa and with an absolute temperature ratio of two. The engine is now under test in a new test facility which was designed and built at NASA Lewis specifically to test the SPRE 1. The SPRE 1, the NASA test facility, the initial SPRE 1 test results, and future SPRE 1 test plans are described.

The primary program goal to requality used bearings by regrinding the races and replacing the balls is briefly reviewed. Southwest Research Institute's task of inspecting the bearings before and after regrinding and endurance testing the refurbished bearings are briefly discussed. Five hundred rejected used T53 engine bearings, part number 1-300-015-(02/04), were inspected utilizing the CIBLE (Critical Inspection of Bearings for Life Extension) system. After the initial inspection was complete, the bearings were shipped to ITI (Industrial Techtonics Incorporated) to be refurbished. The refurbished bearings were returned to SwRI where they were again inspected utilizing the CIBLE system. The results of the CIBLE inspections, before and after regrinding, are presented. Twenty-eight refurnished bearings and 10 new bearings were endurance tested to determine if the refurnished bearings performed as well as new bearings. Conclusions and recommendation, based upon inspections and endurance testing performed at SwRI, are presented. Keywords: Nondestructive evaluation, Bearing inspection, and Endurance testing.

The United States Air Force is currently pursuing a national program to develop the technology necessary to build a single staged aircraft capable of flying from conventional runways into orbit. This vehicle, the National Aerospace Plane (NASP), will use airbreathing propulsion to accelerate through hypersonic speeds to orbital velocities. In order to achieve the aerodynamic and propulsive efficiencies necessary to accomplish this goal, the vehicle forebody must be used as a precompression surface for the engine flowpath. This results in a thick boundary layer which must be bled-off or ingested by the engine. A computational model for thickening boundary layers with mass addition is developed. The phenomena of uniform injection into a two-dimensional supersonic stream and subsequent boundary layer growth downstream is discussed. Analysis of the injection region provides the thickness of the boundary layer just aft of injection. An injection region velocity profile is then used to approximate the boundary layer profile just aft of injection and is input into a finite- difference boundary layer code. Downstream profiles and thicknesses are calculated and compared to experimental results. The computational model developed here provides a tool for the design of a boundary layer generation system for hypersonic engine inlet testing. This mass addition system is needed to simulate the boundary layer developed on the forebody of hypersonic vehicles. Theses. (aw)

Report is a study of the use, timing, and costs of development testing in the new aeronautical test facilities: the Aeropropulsion Systems Test Facility (ASTF), the National Transonic Facility (NTF), and the 80 ft x 120 ft low speed tunnel at NASA-Ames Research Center, California. Keywords: wind tunnels, aircraft engine testing, integrated testing.

Rotating detonation engines (RDEs) have the potential for greater efficiencies over conventional engines by utilizing pressure gain combustion. A new modular RDE (6 in diameter) was developed and successfully run on hydrogen and standard air. The RDE allows for variation of injection scheme and detonation channel widths. Tests provided the operational space of the new RDE as well as characterized detonation unsteadiness. It was found that a smaller equivalence ratio than previous was required to obtain continuous detonations. Also discovered was VCJ was reached in the RDE, but not sustained.

Eleven firings of the Rocketdyne J-2 rocket engine were conducted in Test Cell J-4 of the Large Rocket Facility. The firings were accomplished at pressure altitudes ranging from 85,000 to 109,000 ft at engine start to investigate (1) engine start transients for both S-IVB first burn and (2) 80-min orbital restart with the engine in a 230,000-lbf-thrust configuration. Engine components were thermally conditioned to temperatures representative of an S-IVB first burn or 80-min orbital restart. The total accumulated firing duration for the four test periods was 188.43 sec. Engine operation appeared to be satisfactory.

Ten firings of the Rocketdyne J-2 rocket engine (S/N J-2036-1) were conducted during two test periods (J4-1901-07 and J4-1901-08) on August 21 and 27, 1968, in Test Cell J-4 of the Large Rocket Facility. This testing was in support of the Saturn V vehicle. The firings were accomplished at pressure altitudes between 93,000 and 102,000 ft at engine start. The primary objective of these firings was to evaluate transient gas generator temperatures under simulated engine first burn and restart conditions. Transient gas generator temperatures, particularly on engine restarts were significantly reduced by utilizing a gas generator oxidizer supply line assembly which incorporated a geometrically different bellows section. The supply line assembly utilized on this series of firings had a 1.7 percent greater pressure drop at a flow of 3.5 lbm/sec. The total accumulated firing duration for these tests was 163 sec.

The firings were accomplished at pressure altitudes ranging from 81, 000 to 107,000 ft at engine start. Basic test objectives were to evaluate the effect on the engine start transient of (1) different configuration gas generator oxidizer supply lines, (2) gaseous helium as a start tank pressurant, (3) start tank pressures ranging from 600 to 1000 psia, and (4) a simulated gas generator fuel poppet leak. Engine components were thermally conditioned to temperatures representative of S-IVB engine/interstage environment for a first burn, 80-min restart, or 6-hr restart. The total accumulated firing duration for the three test periods was 181.6 sec.

This report is the product of a 2-year emissions testing program designed to document, characterize, and evaluate emissions from sixteen aircraft engines, two helicopter engines, and two auxiliary power units (APUs) burning JP-8 fuel. The purpose of this engine testing program was to develop emission factors for the tested engines under representative engine load conditions. All testing was performed by the Environmental Quality Management Inc. (EQ) and Roy F. Weston, Inc. (Weston) team. Testing was conducted for criteria pollutants and select hazardous air pollutants (HAPs), e.g., aldehyde/ketones and semivolatile and volatile organic compounds.

The variation in emissions over the exhaust area of a General Electric CF6-50 model engine was investigated in order to determine the requirements for a representative sample. The emission measurements were made in a systematic pattern of 120 sample points using a traversing probe system. These data were used to develop detailed emission profiles at three power levels. At idle power, variations over the exhaust area are attributed to the particular fueling pattern used in current CF6-50 model engines. At higher power levels, where uniform fueling is employed, emission levels are more uniform and are characterized by a slightly peaked radial profile. Average values from the 120-point traverse were comapred with selected 12-point averages in the EPA prescribed cruciform pattern. Generally good agreement between the two averages was obtained. (Author)

The variation in emissions over the exhaust area of a General Electric CF6-50 model engine was investigated in order to determine the requirements for a representative sample. The emission measurements were made in a systematic pattern of 120 sample points using a traversing probe system. These data were used to develop detailed emission profiles at three power levels. At idle power, variations over the exhaust area are attributed to the particular fueling pattern used in current CF6-50 model engines. At higher power levels, where uniform fueling is employed, emission levels are more uniform and are characterized by a slightly peaked radial profile. Average values from the 120-point traverse were comapred with selected 12-point averages in the EPA prescribed cruciform pattern. Generally good agreement between the two averages was obtained. (Author)

Jet engine being set up for testing

This document reports an experimental study of the effectiveness of dry chemical in extinguishing and delaying reignition of fires resulting from hydrocarbon fuel leaking onto heated surfaces such as can occur in jet engine nacelles. The commercial fire extinguishant dry chemicals tried were, for example, sodium and potassium bicarbonate, carbonate, chloride, and carbamate (Monnex) but we have also tested other metal-halogen and metal-hydroxycarbonate compounds prepared in our own laboratories. Given in this report are: synthetic and preparative procedures for new materials developed; a new concept of fire-control by dry chemical agents; descriptions of experimental assemblages to test dry chemical fire extinguishant efficiencies in controlling fuel fires initiated by hot surfaces; comparative testing data for more than 25 chemical systems in a static assemblage with no airflow across the heated surface, and similar comparative data for more than 10 compounds in a dynamic system with airflows up to 350 ft/sec; and recommendations for future work with one system that fulfills all requirements delineated by the sponsoring agency, and which has been tested in both the Static and Dynamic assemblages with both methodologies confirming it as the most effective system by comparison with the other materials tested.

Dynamic data from tests of a T55-L,712 engine are presented. Engine stall/surge data were analyzed using digital signal processing techniques. In addition, forced response testing (system identification studies) was done at various engine speeds. Forced response testing was done using eight jet ejectors approximately equally circumferentially spaced about the compressor front face. This paper presents some preliminary results for the ground idle (approximately 60% of design speed) point. Brief descriptions of the jet injection system, the test matrix, and analysis techniques used are presented. Results of these analyses indicate a substantial transfer of energy across the compressor first stage at some frequencies and that the ejectors are effective in modifying the local flow conditions in front of the first compressor stage.

NASA and industry partners are working towards fabrication process development to reduce costs and schedules associated with manufacturing liquid rocket engine components with the goal of reducing overall mission costs. One such technique being evaluated is powder-bed fusion or selective laser melting (SLM), commonly referred to as additive manufacturing (AM). The NASA Low Cost Upper Stage Propulsion (LCUSP) program was designed to develop processes and material characterization for GRCop-84 (a NASA Glenn Research Center-developed copper, chrome, niobium alloy) commensurate with powder-bed AM, evaluate bimetallic deposition, and complete testing of a full scale combustion chamber. As part of this development, the process has been transferred to industry partners to enable a long-term supply chain of monolithic copper combustion chambers. To advance the processes further and allow for optimization with multiple materials, NASA is also investigating the feasibility of bimetallic AM chambers. In addition to the LCUSP program, NASA has completed a series of development programs and hot-fire tests to demonstrate SLM GRCop-84 and other AM techniques. NASA's efforts include a 4K lbf thrust liquid oxygen/methane (LOX/CH4) combustion chamber and subscale thrust chambers for 1.2K lbf LOX/hydrogen (H2) applications that have been designed and fabricated with SLM GRCop-84. The same technologies for these lower thrust applications are being applied to 25-35K lbf main combustion chamber (MCC) designs. This paper describes the design, development, manufacturing and testing of these numerous combustion chambers, and the associated lessons learned throughout their design and development processes.

NASA and industry partners are working towards fabrication process development to reduce costs and schedules associated with manufacturing liquid rocket engine components with the goal of reducing overall mission costs. One such technique being evaluated is powder-bed fusion or selective laser melting (SLM), commonly referred to as additive manufacturing (AM). The NASA Low Cost Upper Stage Propulsion (LCUSP) program was designed to develop processes and material characterization for GRCop-84 (a NASA Glenn Research Center-developed copper, chrome, niobium alloy) commensurate with powder-bed AM, evaluate bimetallic deposition, and complete testing of a full scale combustion chamber. As part of this development, the process has been transferred to industry partners to enable a long-term supply chain of monolithic copper combustion chambers. To advance the processes further and allow for optimization with multiple materials, NASA is also investigating the feasibility of bimetallic AM chambers. In addition to the LCUSP program, NASA has completed a series of development programs and hot-fire tests to demonstrate SLM GRCop-84 and other AM techniques. NASA's efforts include a 4K lbf thrust liquid oxygen/methane (LOX/CH4) combustion chamber and subscale thrust chambers for 1.2K lbf LOX/hydrogen (H2) applications that have been designed and fabricated with SLM GRCop-84. The same technologies for these lower thrust applications are being applied to 25-35K lbf main combustion chamber (MCC) designs. This paper describes the design, development, manufacturing and testing of these numerous combustion chambers, and the associated lessons learned throughout their design and development processes.

Five firings of the J-2S rocket engine were conducted in Rocket Development Test Cell (J-4) on July 17, and 29, 1969. These firings were accomplished at pressure altitudes ranging from 85,000 to 101,000 ft at engine start. The primary objectives of these test periods were to (1) determine if main-stage conditions which existed during sea-level testing of engine S/N J-113 would result in similar abnormal oxidizer dome vibrations in the 4400- to 4700- Hz frequency range during altitude testing of engine J-112-E, (2) evaluate high thrust idle-mode operation with a simulated full-face oxidizer flow injector configuration, and (3) document effects of closing the thrust chamber bypass valve during high thrust idle-mode operation. Altitude testing did not result in abnormal (greater than 100 grams) oxidizer dome vibration in the 4400- to 4700- Hz range during test period 04. The thrust chamber bypass valve closing resulted in a 65 F increase in fuel injection temperature; however, stabilized high thrust idle-mode operation was not attained during test period 05.

Five firings of the J-2S rocket engine were conducted in Rocket Development Test Cell (J-4) on July 17, and 29, 1969. These firings were accomplished at pressure altitudes ranging from 85,000 to 101,000 ft at engine start. The primary objectives of these test periods were to (1) determine if main-stage conditions which existed during sea-level testing of engine S/N J-113 would result in similar abnormal oxidizer dome vibrations in the 4400- to 4700- Hz frequency range during altitude testing of engine J-112-E, (2) evaluate high thrust idle-mode operation with a simulated full-face oxidizer flow injector configuration, and (3) document effects of closing the thrust chamber bypass valve during high thrust idle-mode operation. Altitude testing did not result in abnormal (greater than 100 grams) oxidizer dome vibration in the 4400- to 4700- Hz range during test period 04. The thrust chamber bypass valve closing resulted in a 65 F increase in fuel injection temperature; however, stabilized high thrust idle-mode operation was not attained during test period 05.

Two builds of an F100 engine model derivative (EMD) engine were evaluated for improvements in engine components and digital electronic engine control (DEEC) logic. Two DEEC flight logics were verified throughout the flight envelope in support of flight clearance for the F100 engine model derivative program (EMPD). A nozzle instability and a faster augmentor transient capability was investigated in support of the F-15 DEEC flight program. Off schedule coupled system mode fan flutter, DEEC nose-boom pressure correlation, DEEC station six pressure comparison, and a new fan inlet variable vane (CIVV) schedule are identified.

Since 1984, an effort has been underway at Rocketdyne, manufacturer of the Space Shuttle Main Engine (SSME), to automate much of the analysis procedure conducted after engine test firings. Previously published articles at national and international conferences have contained the context of and justification for this effort. Here, progress is reported in building the full system, including the extensions of integrating large databases with the system, known as Scotty. Inductive knowledge acquisition has proven itself to be a key factor in the success of Scotty. The combination of a powerful inductive expert system building tool (ExTran), a relational data base management system (Reliance), and software engineering principles and Computer-Assisted Software Engineering (CASE) tools makes for a practical, useful and state-of-the-art application of an expert system.

A trade study of the feasibility of conducting J-2X testing in the Glenn Research Center (GRC) Plum Brook Station (PBS) B-2 facility was initiated in May 2006 with results available in October 2006. The Propulsion Test Integration Group (PTIG) led the study with support from Marshall Space Flight Center (MSFC) and Jacobs Sverdrup Engineering. The primary focus of the trade study was on facility design concepts and their capability to satisfy the J-2X altitude simulation test requirements. The propulsion systems tested in the B-2 facility were in the 30,000-pound (30K) thrust class. The J-2X thrust is approximately 10 times larger. Therefore, concepts significantly different from the current configuration are necessary for the diffuser, spray chamber subsystems, and cooling water. Steam exhaust condensation in the spray chamber is judged to be the key risk consideration relative to acceptable spray chamber pressure. Further assessment via computational fluid dynamics (CFD) and other simulation capabilities (e.g. methodology for anchoring predictions with actual test data and subscale testing to support investigation.

Mission: Common upper stage engine for Ares I and Ares V. Challenge: Use proven technology from Saturn X-33, RS-68 to develop the highest Isp GG cycle engine in history for 2 missions in record time . Key Features: LOX/LH2 GG cycle, series turbines (2), HIP-bonded MCC, pneumatic ball-sector valves, on-board engine controller, tube-wall regen nozzle/large passively-cooled nozzle extension, TEG boost/cooling . Development Philosophy: proven hardware, aggressive schedule, early risk reduction, requirements-driven.

The design of a subscale jet engine driven ejector/diffuser system is examined. Analytical results and preliminary design drawings and plans are included. Previously developed performance prediction techniques are verified. A safety analysis is performed to determine the mechanism for detonation suppression.