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NASA NTRS1993

STS-55 pad abort: Engine 2011 oxidizer preburner augmented spark igniter check valve leak

The STS-55 initial launch attempt of Columbia (OV102) was terminated on KSC launch pad A March 22, 1993 at 9:51 AM E.S.T. due to violation of an ME-3 (Engine 2011) Launch Commit Criteria (LCC) limit exceedance. The event description and timeline are summarized. Propellant loading was initiated on 22 March, 1993 at 1:15 AM EST. All SSME chill parameters and launch commit criteria (LCC) were nominal. At engine start plus 1.44 seconds, a Failure Identification (FID) was posted against Engine 2011 for exceeding the 50 psia Oxidizer Preburner (OPB) purge pressure redline. The engine was shut down at 1.50 seconds followed by Engines 2034 and 2030. All shut down sequences were nominal and the mission was safely aborted. The OPB purge pressure redline violation and the abort profile/overlay for all three engines are depicted. SSME Avionics hardware and software performed nominally during the incident. A review of vehicle data table (VDT) data and controller software logic revealed no failure indications other than the single FID 013-414, OPB purge pressure redline exceeded. Software logic was executed according to requirements and there was no anomalous controller software operation. Immediately following the abort, a Rocketdyne/NASA failure investigation team was assembled. The team successfully isolated the failure cause to the oxidizer preburner augmented spark igniter purge check valve not being fully closed due to contamination. The source of the contaminant was traced to a cut segment from a rubber O-ring which was used in a fine clean tool during valve production prior to 1992. The valve was apparently contaminated during its fabrication in 1985. The valve had performed acceptably on four previous flights of the engine, and SSME flight history shows 780 combined check valve flights without failure. The failure of an Engine 3 (SSME No. 2011) check valve to close was sensed by onboard engine instruments even though all other engine operations were normal. This resulted in an engine shutdown and safe sequential shutdown of all three engines prior to ignition of the solid boosters.

Source record↗
NASA NTRS1993

STS-51 pad abort. OV103-engine 2033 (ME-2) fuel flowmeter sensor open circuit

The STS-51 initial launch attempt of Discovery (OV-103) was terminated on KSC launch pad 39B on 12 Aug. 1993 at 9:12 AM E.S.T. due to a sensor redundancy failure in the liquid hydrogen system of ME-2 (Engine 2033). The event description and time line are summarized. Propellant loading was initiated on 12 Aug. 1993 at 12:00 AM EST. All space shuttle main engine (SSME) chill parameters and Launch Commit Criteria (LCC) were nominal. At engine start plus 1.34 seconds a Failure Identification (FID) was posted against Engine 2033 for exceeding the 1800 spin intra-channel (A1-A2) Fuel Flowrate sensor channel qualification limit. The engine was shut down at 1.50 seconds followed by Engines 2032 and 2030. All shut down sequences were nominal and the mission was safely aborted. SSME Avionics hardware and software performed nominally during the incident. A review of vehicle data table (VDT) data and controller software logic revealed no failure indications other than the single FID 111-101, Fuel Flowrate Intra-Channel Test Channel A disqualification. Software logic was executed according to requirements and there was no anomalous controller software operation. Immediately following the abort, a Rocketdyne/NASA failure investigation team was assembled. The team successfully isolated the failure cause to an open circuit in a Fuel Flowrate Sensor. This type of failure has occurred eight previous times in ground testing. The sensor had performed acceptably on three previous flights of the engine and SSME flight history shows 684 combined fuel flow rate sensor channel flights without failure. The disqualification of an Engine 2 (SSME No. 2033) Fuel Flowrate sensor channel was a result of an instrumentation failure and not engine performance. All other engine operations were nominal. This disqualification resulted in an engine shutdown and safe sequential shutdown of all three engines prior to ignition of the solid boosters.

Source record↗
NASA NTRS2002

Modes 1 and 4 Testing of an Axisymmetric GTX RBCC Engine

Under sponsorship of the NASA Glenn Research Center (NASA GRC), the Johns Hopkins University Applied Physics Laboratory (JHU/APL) designed and built a five-inch diameter, Rocket-Based Combined Cycle (RBCC) engine to investigate mode 1 and mode 4 performance. This engine was designed so that engine area and length ratios were similar to the NASA GRC GTX engine. Unlike the GTX semi-circular engine design, the APL engine is completely axisymmetric. For this design, a traditional rocket thruster was installed inside of the scramjet flowpath, along the engine centerline. A three part test series is currently underway to determine mode 1 and mode 4 engine performance and compare those results to the rocket thruster operating by itself. All tests are to be conducted in APL's Cell 2 Freejet Engine Facility. Testing of the rocket thruster alone has been accomplished and its performance determined (average Isp efficiency = 90%). Mode 1 (air-augmented rocket) testing is to be conducted at various chamber pressure-to-ambient pressure ratios with the engine inlet fully open. Mode 4 (vacuum rocket) testing is to be conducted with chamber pressure-to-ambient pressure ratios greater than 600 with the inlet fully closed to simulate high-altitude flight conditions. The engine is currently installed in Cell 2 and awaiting testing. This paper discusses the engine design and details the results of the rocket-only testing. Comparisons to previous tests and projected mode 1 and mode 4 test conditions are also discussed.

Rice, Tharen↗
NASA NTRS2002

Liquid Engine Test Facilities Assessment

The John C. Stennis Space Center (SSC) requested The Aerospace Corporation to examine the current testing capability of all existing large liquid engine test facilities located in the United States. That information along with projected liquid rocket engine development was used to examine future liquid rocket engine testing facilities needs in the coming decade. Current domestic liquid engine test facilities capabilities, when examined against engine concepts for the coming decade, indicate there are ample facilities offering altitude simulation during test. In addition, it was observed that many contractor facilities have limited ambient test capability of larger thrust engines under current consideration. Finally, it was concluded that diminished contractor participation engine development testing will drive this activity to the government sector. Only three facilities are seen as key contributors to engine testing in the coming decade, namely John C. Stennis Space Center (SSC), Marshall Space Flight Center (MSFC), and Air Force Research Laboratory (AFRL). Past rocket engine test experience was evaluated as a possible resource for projecting future engine test needs. A database comprised of various engine models and the level of testing performed to flight qualify those systems for their first flight was constructed. For comparison purposes in this study, development and qualification efforts were totaled and treated as one test program. Based on experience with past Air Force programs, the time on the test stand accounts for typically 50% or more of the total program time. Historical data show that the time to design and develop new engines has increased over the last 40 years, most likely due to scarcer resources in today's funding environment.

Adams, Michael J.↗
NASA NTRS2003

Data-Mining Toolset Developed for Determining Turbine Engine Part Life Consumption

The current practice in aerospace turbine engine maintenance is to remove components defined as life-limited parts after a fixed time, on the basis of a predetermined number of flight cycles. Under this schedule-based maintenance practice, the worst-case usage scenario is used to determine the usable life of the component. As shown, this practice often requires removing a part before its useful life is fully consumed, thus leading to higher maintenance cost. To address this issue, the NASA Glenn Research Center, in a collaborative effort with Pratt & Whitney, has developed a generic modular toolset that uses data-mining technology to parameterize life usage models for maintenance purposes. The toolset enables a "condition-based" maintenance approach, where parts are removed on the basis of the cumulative history of the severity of operation they have experienced. The toolset uses data-mining technology to tune life-consumption models on the basis of operating and maintenance histories. The flight operating conditions, represented by measured variables within the engine, are correlated with repair records for the engines, generating a relationship between the operating condition of the part and its service life. As shown, with the condition-based maintenance approach, the lifelimited part is in service until its usable life is fully consumed. This approach will lower maintenance costs while maintaining the safety of the propulsion system. The toolset is a modular program that is easily customizable by users. First, appropriate parametric damage accumulation models, which will be functions of engine variables, must be defined. The tool then optimizes the models to match the historical data by computing an effective-cycle metric that reduces the unexplained variability in component life due to each damage mode by accounting for the variability in operational severity. The damage increment due to operating conditions experienced during each flight is used to compute the effective cycles and ultimately the replacement time. Utilities to handle data problems, such as gaps in the flight data records, are included in the toolset. The tool was demonstrated using the first stage, high-pressure turbine blade of the PW4077 engine (Pratt & Whitney, East Hartford, CT). The damage modes considered were thermomechanical fatigue and oxidation/erosion. Each PW4077 engine contains 82 first-stage, high-pressure turbine blades, and data from a fleet of engines were used to tune the life-consumption models. The models took into account not only measured variables within the engine, but also unmeasured variables such as engine health parameters that are affected by degradation of the engine due to aging. The tool proved effective at predicting the average number of blades scrapped over time due to each damage mode, per engine, given the operating history of the engine. The customizable tools are available to interested parties within the aerospace community.

Litt, Jonathan S.↗
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 NTRS2011

From Paper to Production to Test: An Update on NASA's J-2X Engine for Exploration

The NASA/industry team responsible for developing the J-2X upper stage engine for the Space Launch System (SLS) Program has made significant progress toward moving beyond the design phase and into production, assembly, and test of development hardware. The J-2X engine exemplifies the SLS Program goal of using proven technology and experience from more than 50 years of United States spaceflight experience combined with modern manufacturing processes and approaches. It will power the second stage of the fully evolved SLS Program launch vehicle that will enable a return to human exploration of space beyond low earth orbit. Pratt & Whitney Rocketdyne (PWR) is under contract to develop and produce the engine, leveraging its flight-proven LH2/LOX, gas generator cycle J-2 and RS-68 engine capabilities, recent experience with the X-33 aerospike XRS-2200 engine, and development knowledge of the J-2S tap-off cycle engine. The J- 2X employs a gas generator operating cycle designed to produce 294,000 pounds of vacuum thrust in primary operating mode with its full nozzle extension. With a truncated nozzle extension suitable to support engine clustering on the stage, the nominal vacuum thrust level in primary mode is 285,000 pounds. It also has a secondary mode, during which it operates at 80 percent thrust by altering its mixture ratio. The J-2X development philosophy is based on proven hardware, an aggressive development schedule, and early risk reduction. NASA Marshall Space Flight Center (MSFC) and PWR began development of the J-2X in June 2006. The government/industry team of more than 600 people within NASA and PWR successfully completed the Critical Design Review (CDR) in November 2008, following extensive risk mitigation testing. Assembly of the first development engine was completed in May 2011 and the first engine test was conducted at the NASA Stennis Space Center (SSC), test stand A2, on 14 July 2011. Testing of the first development engine will continue through the autumn of 2011, be paused for test stand modifications to the passive diffuser, and then restart in the spring of 2012. This testing will be followed by specialized powerpack testing intended to examine the design and operating margins of the engine turbomachinery. The development plan beyond this point leads through more system-level, engine testing of several samples, analytical model validation activities, functional and performance verification, and then ultimate certification to support human spaceflight. This paper will discuss the J-2X development background, provide top-level information on design and development planning, and will explore some of the development challenges and mitigation activities pursued to date.

Kynard, Michael↗
NASA NTRS2013

Distributed Engine Control Empirical/Analytical Verification Tools

NASA's vision for an intelligent engine will be realized with the development of a truly distributed control system featuring highly reliable, modular, and dependable components capable of both surviving the harsh engine operating environment and decentralized functionality. A set of control system verification tools was developed and applied to a C-MAPSS40K engine model, and metrics were established to assess the stability and performance of these control systems on the same platform. A software tool was developed that allows designers to assemble easily a distributed control system in software and immediately assess the overall impacts of the system on the target (simulated) platform, allowing control system designers to converge rapidly on acceptable architectures with consideration to all required hardware elements. The software developed in this program will be installed on a distributed hardware-in-the-loop (DHIL) simulation tool to assist NASA and the Distributed Engine Control Working Group (DECWG) in integrating DCS (distributed engine control systems) components onto existing and next-generation engines.The distributed engine control simulator blockset for MATLAB/Simulink and hardware simulator provides the capability to simulate virtual subcomponents, as well as swap actual subcomponents for hardware-in-the-loop (HIL) analysis. Subcomponents can be the communication network, smart sensor or actuator nodes, or a centralized control system. The distributed engine control blockset for MATLAB/Simulink is a software development tool. The software includes an engine simulation, a communication network simulation, control algorithms, and analysis algorithms set up in a modular environment for rapid simulation of different network architectures; the hardware consists of an embedded device running parts of the CMAPSS engine simulator and controlled through Simulink. The distributed engine control simulation, evaluation, and analysis technology provides unique capabilities to study the effects of a given change to the control system in the context of the distributed paradigm. The simulation tool can support treatment of all components within the control system, both virtual and real; these include communication data network, smart sensor and actuator nodes, centralized control system (FADEC full authority digital engine control), and the aircraft engine itself. The DECsim tool can allow simulation-based prototyping of control laws, control architectures, and decentralization strategies before hardware is integrated into the system. With the configuration specified, the simulator allows a variety of key factors to be systematically assessed. Such factors include control system performance, reliability, weight, and bandwidth utilization.

DeCastro, Jonathan↗
DOE OSTI2024

Recovery of Natural Gas Equipment Emissions into Gas Compression Engines for the Reduction of Potential Greenhouse Gas Emissions

Since the turn of the millennium, the United States (U.S.) oil and natural gas (ONG) industry has nearly doubled its natural gas production rate. As a result, the ONG industry has recently come under increasing scrutiny for its contributions to greenhouse gas (GHG) emissions. Consequently, various solutions to this problem have been proposed and formulated to reduce the impacts of GHG emissions on the environment. West Virginia University (WVU) have found it important to research the impacts of recovering vented gas streams into prime-mover engines. The U.S. Department of Energy (DOE) and National Energy Technology Laboratory (NETL) have granted WVU funding to research and develop a “Methane Mitigator” (M2) - a “Scalable Vent Mitigation Strategy to Simultaneously Reduce Methane Emissions and Fuel Consumption from the Compression Industry.” One of the main areas of interest for this research was the collection of emissions from natural gas equipment into a Caterpillar G3508J natural gas compression engine. The parameters being analyzed from the engine were brake-specific emissions and power output. The emissions sources considered for this research were pneumatic controllers (PCs), reciprocating compressor vents, and the engine’s open crankcase breather. The compressor vent and PC emissions were simulated using a mass flow controller (MFC) and flowed into the engine using two separate methods: (1) directly into the air intake, and (2) through a retrofitted closed crankcase ventilation system (CCV), serving as a buffer volume. The crankcase emissions were quantified without the CCV, and the impact on exhaust emissions from circulating the crankcase gases into the intake was measured. The simulated compressor vent and PC flows from the MFC had limited effect on the steady state operation of the engine and resulting performance. When the simulated flows were fed directly into the engine’s air intake, the changes within the engine’s continuous performance and emission parameters were larger but lasted for shorter durations. Conversely, when the simulated flows were fed into the CCV before entering the air intake, the changes in the engine’s performance and emission parameters were less pronounced for continuous analysis but lasted for longer durations. In either case, the continuous emission changes in both emissions and performance varied in size depending on the test scenario being run, but the cycle average changes in emissions and performance showed little impact overall compared to the engine’s baseline operation. As a result, the inclusion of a CCV shows a decrease in baseline carbon dioxide equivalent (CO2-eq.) engine emissions (from combined exhaust and open crankcase) of almost 4%. Likewise, the CCV inclusion reduced baseline total methane (CH4) from combined exhaust and open crankcase by upwards of 16%. These atmospheric emissions only decreased further with the inclusions of collected PC and compressor vent flows. The resulting changes in time-averaged rated exhaust behavior (or lack thereof) prove that the proposed M2 system could likely be deployed at sites with modern lean-burn natural gas engines as a viable option for reducing and eliminating potential GHG sources that would have otherwise been unutilized as energy sources.

03 NATURAL GAS↗
NASA NTRS2015

A Fully Non-Metallic Gas Turbine Engine Enabled by Additive Manufacturing Part I: System Analysis, Component Identification, Additive Manufacturing, and Testing of Polymer Composites

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.

acoustic liners↗
NASA NTRS2006

Advancing the practice of systems engineering at JPL

In FY 2004, JPL launched an initiative to improve the way it practices systems engineering. The Lab's senior management formed the Systems Engineering Advancement (SEA) Project in order to "significantly advance the practice and organizational capabilities of systems engineering at JPL on flight projects and ground support tasks." The scope of the SEA Project includes the systems engineering work performed in all three dimensions of a program, project, or task: 1. the full life-cycle, i.e., concept through end of operations 2. the full depth, i.e., Program, Project, System, Subsystem, Element (SE Levels 1 to 5) 3. the full technical scope, e.g., the flight, ground and launch systems, avionics, power, propulsion, telecommunications, thermal, etc. The initial focus of their efforts defined the following basic systems engineering functions at JPL: systems architecture, requirements management, interface definition, technical resource management, system design and analysis, system verification and validation, risk management, technical peer reviews, design process management and systems engineering task management, They also developed a list of highly valued personal behaviors of systems engineers, and are working to inculcate those behaviors into members of their systems engineering community. The SEA Project is developing products, services, and training to support managers and practitioners throughout the entire system lifecycle. As these are developed, each one needs to be systematically deployed. Hence, the SEA Project developed a deployment process that includes four aspects: infrastructure and operations, communication and outreach, education and training, and consulting support. In addition, the SEA Project has taken a proactive approach to organizational change management and customer relationship management - both concepts and approaches not usually invoked in an engineering environment. This paper'3 describes JPL's approach to advancing the practice of systems engineering at the Lab. It describes the general approach used and how they addressed the three key aspects of change: people, process and technology. It highlights a list of highly valued personal behaviors of systems engineers, discusses the various products, services and training that were developed, describes the deployment approach used, and concludes with several lessons learned.

process improvement↗
NASA NTRS1979

Orbital transfer vehicle advanced expander cycle engine point design study

A detailed computer model of the expander cycle engine system steady-state operation was developed. Payload performance data for the expander cycle engine for 100K pounds and 65K pounds Space Shuttle missions were examined to establish peak payload thrusts. It was found that peak payload for the 100K pound gross weight single engine case occurred at nearly 15,000 pounds thrust for both engine mixture ratios of 6:1 and 7:1. Peak payload at engine mixture ratio of 6:1 was 4 percent greater than at 7:1 mixture ratio. The two engine configuration peak paYloads occurred at engine thrust below 7,500 pounds (extrapolated) and were lower than single engine payloads. For the 65K pound gross weight mission, peak payloads for the single engine configuration occurred at nearly 10,000 pounds for both mixture ratios. Mixture ratio of 7:1 payloads were 5 percent lower than 6:1 payloads. Two engine configuration payloads occurred at engine thrusts considerably below 10K pounds.

Martinez, A.↗
NASA NTRS1981

Apparatus for sensor failure detection and correction in a gas turbine engine control system

A gas turbine engine control system maintains a selected level of engine performance despite the failure or abnormal operation of one or more engine parameter sensors. The control system employs a continuously updated engine model which simulates engine performance and generates signals representing real time estimates of the engine parameter sensor signals. The estimate signals are transmitted to a control computational unit which utilizes them in lieu of the actual engine parameter sensor signals to control the operation of the engine. The estimate signals are also compared with the corresponding actual engine parameter sensor signals and the resulting difference signals are utilized to update the engine model. If a particular difference signal exceeds specific tolerance limits, the difference signal is inhibited from updating the model and a sensor failure indication is provided to the engine operator.

Spang, H. A., III↗
NASA NTRS1983

Stirling engine application study

A range of potential applications for Stirling engines in the power range from 0.5 to 5000 hp is surveyed. Over one hundred such engine applications are grouped into a small number of classes (10), with the application in each class having a high degree of commonality in technical performance and cost requirements. A review of conventional engines (usually spark ignition or Diesel) was then undertaken to determine the degree to which commercial engine practice now serves the needs of the application classes and to detemine the nature of the competition faced by a new engine system. In each application class the Stirling engine was compared to the conventional engines, assuming that objectives of ongoing Stirling engine development programs are met. This ranking process indicated that Stirling engines showed potential for use in all application classes except very light duty applications (lawn mowers, etc.). However, this potential is contingent on demonstrating much greater operating life and reliability than has been demonstrated to date by developmental Stirling engine systems. This implies that future program initiatives in developing Stirling engine systems should give more emphasis to life and reliability issues than has been the case in ongoing programs.

Teagan, W. P.↗
NASA NTRS1983

Multi-Fuel Rotary Engine for General Aviation Aircraft

Design studies, conducted for NASA, of Advanced Multi-fuel General Aviation and Commuter Aircraft Rotary Stratified Charge Engines are summarized. Conceptual design studies of an advanced engine sized to provide 186/250 shaft KW/HP under cruise conditions at 7620/25,000 m/ft. altitude were performed. Relevant engine development background covering both prior and recent engine test results of the direct injected unthrottled rotary engine technology, including the capability to interchangeably operate on gasoline, diesel fuel, kerosene, or aviation jet fuel, are presented and related to growth predictions. Aircraft studies, using these resultant growth engines, define anticipated system effects of the performance and power density improvements for both single engine and twin engine airplanes. The calculated results indicate superior system performance and 30 to 35% fuel economy improvement for the Rotary-engine airplanes as compared to equivalent airframe concept designs with current baseline engines. The research and technology activities required to attain the projected engine performance levels are also discussed.

Jones, C.↗
NASA NTRS1985

Harrier III-AV8B with a modern engine

The paper examines the application of a 26, 156-lb thrust engine (which is 30 percent more thrust than the short-lift dry rating on the F402-RR-406 Pegasus engine) to an AV-8B aircraft. This modern engine has 30 percent more thrust than the F402-RR-406 Pegasus engine with no increase in engine weight; the weight is assumed to be constant for both engines. The paper focuses on the benefit to the AV-8B in increasing short takeoff gross weight and vertical landing weight. A second comparison is made between the bigger engine with the existing thrust-versus-SFC curve, and a 10 percent lower SFC curve. The notion of a modern engine with 10 percent lower SFC and 30 percent more thrust for the same size and weight is presented as a hypothesis, and is not supported by discussing engine technology. There are no engine data or technical discussion as to how the modern engine is constructed or which manufacturer could build it. The performance estimates are generated by ACSYNT, a NASA developed computer code that is available to U.S. industry and has been reported on in a number of other publications.

Wilson, S. B., III↗
NASA NTRS1989

A supersonic through-flow fan engine airframe integration study

A study is undertaken to investigate the engine airframe integration effects for supersonic through-flow fan engines installed on a Mach 3.20 supersonic cruise vehicle. Six different supersonic through-flow fan engine installations covering the effects of engine size, nacelle contour, nacelle placement, and approximate bypass plume effects are presented. The different supersonic through-flow fan installations are compared with a conventional turbine bypass engine configuration on the same basic airframe. The supersonic through-flow fan engine integrations are shown to be comparable to the turbine bypass engine configuration on the basis of installed nacelle wave drag. The supersonic through-flow fan engine airframe integrated vehicles have superior aerodynamic performance on the basis of maximum lift-to-drag ratio than the turbine bypass engine installation over the entire operating Mach number range from 1.10 to 3.20. When approximate bypass plume modeling is included, the supersonic through-flow fan engine configuration shows even larger improvements over the turbine bypass engine configuration.

Barnhart, Paul J.↗
NASA NTRS1989

A supersonic through-flow fan engine airframe integration study

Engine airframe integration effects are investigated for supersonic through-flow fan engines installed on a Mach 3.20 supersonic cruise vehicle. Six different supersonic through-flow fan engine installations covering the effects of engine size, nacelle contour, nacelle placement, and approximate bypass plume effects are presented. The different supersonic through-flow fan installations are compared with a conventional turbine bypass engine configuration on the same basic airframe. The supersonic through-flow fan engine integrations are shown to be comparable to the turbine bypass engine configuration on the basis of installed nacelle wave drag. The supersonic through-flow fan engine airframe integrated vehicles have superior aerodynamic performance on the basis of maximum lift-to-drag ratio than the turbine bypass engine installation over the entire operating Mach number range from 1.10 to 3.20. When approximate bypass plume modeling is included, the supersonic through-flow fan engine configuration shows even larger improvements over the turbine bypass engine configuration.

Barnhart, Paul J.↗