SEARCH · Search NASA
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
The turbulent diffusion of a hot air jet in air can be divided into two zones, the core and the transition zone. The present report deals with the transitional zone.
As part of NASA's program to develop technology for short takeoff and vertical landing (STOVL) fighter aircraft, control system designs have been developed for a conceptual STOVL aircraft. This aircraft is representative of the class of mixed-flow remote-lift concepts that was identified as the preferred design approach by the U.S./U.K. STOVL Joint Assessment and Ranking Team. The control system designs have been evaluated throughout the powered-lift flight envelope on the Vertical Motion Simulator (VMS) at Ames Research Center. Items assessed in the control system evaluation were: maximum control power used in transition and vertical flight, control system dynamic response associated with thrust transfer for attitude control, thrust margin in the presence of ground effect and hot-gas ingestion, and dynamic thrust response for the engine core. Effects of wind, turbulence, and ship airwake disturbances are incorporated in the evaluation. Results provide the basis for a reassessment of existing flying-qualities design criteria applied to STOVL aircraft.
An analysis of gas turbine engines using water and oxygen injection to enhance performance by increasing Mach number capability and by increasing thrust is described. The liquids are injected, either separately or together, into the subsonic diffuser ahead of the engine compressor. A turbojet engine and a mixed-flow turbofan engine (MFTF) are examined, and in pursuit of maximum thrust, both engines are fitted with afterburners. The results indicate that water injection alone can extend the performance envelope of both engine types by one and one-half Mach numbers at which point water-air ratios reach 17 or 18 percent and liquid specific impulse is reduced to some 390 to 470 seconds, a level about equal to the impulse of a high energy rocket engine. The envelope can be further extended, but only with increasing sacrifices in liquid specific impulse. Oxygen-airflow ratios as high as 15 percent were investigated for increasing thrust. Using 15 percent oxygen in combination with water injection at high supersonic Mach numbers resulted in thrust augmentation as high as 76 percent without any significant decrease in liquid specific impulse. The stoichiometric afterburner exit temperature increased with increasing oxygen flow, reaching 4822 deg R in the turbojet engine at a Mach number of 3.5. At the transonic Mach number of 0.95 where no water injection is needed, an oxygen-air ratio of 15 percent increased thrust by some 55 percent in both engines, along with a decrease in liquid specific impulse of 62 percent. Afterburner temperature was approximately 4700 deg R at this high thrust condition. Water and/or oxygen injection are simple and straightforward strategies to improve engine performance and they will add little to engine weight. However, if large Mach number and thrust increases are required, liquid flows become significant, so that operation at these conditions will necessarily be of short duration.
As part of NASA's program to develop technology for short takeoff and vertical landing (STOVL) fighter aircraft, control system designs have been developed for a conceptual STOVL aircraft. This aircraft is representative of the class of mixed-flow remote-lift concepts that was identified as the preferred design approach by the US/UK STOVL Joint Assessment and Ranking Team. The control system designs have been evaluated throughout the powered-lift flight envelope on Ames Research Center's Vertical Motion Simulator. Items assessed in the control system evaluation were: maximum control power used in transition and vertical flight, control system dynamic response associated with thrust transfer for attitude control, thrust margin in the presence of ground effect and hot gas ingestion, and dynamic thrust response for the engine core. Effects of wind, turbulence, and ship airwake disturbances are incorporated in the evaluation. Results provide the basis for a reassessment of existing flying qualities design criteria applied to STOVL aircraft.
A numerical investigation of exergy destroyed by operation of a conventional steam power plant is computed via an exergy cascade. An order of magnitude analysis shows that exergy destruction is dominated by combustion and heat transfer across temperature differences inside the boiler, and conversion of energy entering the turbine/generator sets from thermal to electrical. Combustion and heat transfer inside the boiler accounts for 53.83 percent of the total exergy destruction. Converting thermal energy into electrical energy is responsible for 41.34 percent of the total exergy destruction. Heat transfer across the condenser accounts for 2.89 percent of the total exergy destruction. Fluid flow with friction is responsible for 0.50 percent of the total exergy destruction. The boiler feed pump turbine accounts for 0.25 percent of the total exergy destruction. Fluid flow mixing is responsible for 0.23 percent of the total exergy destruction. Other equipment including gland steam condenser, drain cooler, deaerator and heat exchangers are, in the aggregate, responsible for less than one percent of the total exergy destruction. An energy analysis is also given for comparison of exergy cascade to energy cascade. Efficiencies based on both the first law and second law of thermodynamics are calculated for a number of components and for the plant. The results show that high first law efficiency does not mean high second law efficiency. Therefore, the second law analysis has been proven to be a more powerful tool in pinpointing real losses. The procedure used to determine total exergy destruction and second law efficiency can be used in a conceptual design and parametric study to evaluate the performance of other steam power plants and other thermal systems.
A method for conceptual aircraft design that incorporates the optimization of major engine design variables for a variety of cycle types was developed. The methodology should improve the lengthy screening process currently involved in selecting an appropriate engine cycle for a given application or mission. The new capability will allow environmental concerns such as airport noise and emissions to be addressed early in the design process. The ability to rapidly perform optimization and parametric variations using both engine cycle and aircraft design variables, and to see the impact on the aircraft, should provide insight and guidance for more detailed studies. A brief description of the aircraft performance and mission analysis program and the engine cycle analysis program that were used is given. A new method of predicting propulsion system weight and dimensions using thermodynamic cycle data, preliminary design, and semi-empirical techniques is introduced. Propulsion system performance and weights data generated by the program are compared with industry data and data generated using well established codes. The ability of the optimization techniques to locate an optimum is demonstrated and some of the problems that had to be solved to accomplish this are illustrated. Results from the application of the program to the analysis of three supersonic transport concepts installed with mixed flow turbofans are presented. The results from the application to a Mach 2.4, 5000 n.mi. transport indicate that the optimum bypass ratio is near 0.45 with less than 1 percent variation in minimum gross weight for bypass ratios ranging from 0.3 to 0.6. In the final application of the program, a low sonic boom fix a takeoff gross weight concept that would fly at Mach 2.0 overwater and at Mach 1.6 overland is compared with a baseline concept of the same takeoff gross weight that would fly Mach 2.4 overwater and subsonically overland. The results indicate that for the design mission, the low boom concept has a 5 percent total range penalty relative to the baseline. Additional cycles were optimized for various design overland distances and the effect of flying off-design overland distances is illustrated.
Using a generalized simulation model, a moving-base simulation of a lift-fan short takeoff/vertical landing fighter aircraft was conducted on the Vertical Motion Simulator at Ames Research Center. Objectives of the experiment were to (1) assess the effects of lift-fan propulsion system design features on aircraft control during transition and vertical flight including integration of lift fan/lift/cruise engine/aerodynamic controls and lift fan/lift/cruise engine dynamic response, (2) evaluate pilot-vehicle interface with the control system and head-up display including control modes for low-speed operational tasks and control mode/display integration, and (3) conduct operational evaluations of this configuration during takeoff, transition, and landing similar to those carried out previously by the Ames team for the mixed-flow, vectored thrust, and augmentor-ejector concepts. Based on results of the simulation, preliminary assessments of acceptable and borderline lift-fan and lift/cruise engine thrust response characteristics were obtained. Maximum pitch, roll, and yaw control power used during transition, hover, and vertical landing were documented. Control and display mode options were assessed for their compatibility with a range of land-based and shipboard operations from takeoff to cruise through transition back to hover and vertical landing. Flying qualities were established for candidate control modes and displays for instrument approaches and vertical landings aboard an LPH assault ship and DD-963 destroyer. Test pilot and engineer teams from the Naval Air Warfare Center, Boeing, Lockheed, McDonnell Douglas, and the British Defence Research Agency participated in the program.
An expanded and improved version of a previously published benzene oxidation mechanism is presented and shown to model published experimental data fairly successfully. This benzene submodel is coupled to a modified version of a toluene oxidation submodel from the recent literature. This complete mechanism is shown to successfully model published experimental toluene oxidation data for a highly mixed flow reactor and for higher temperature ignition delay times in a shock tube. A comprehensive sensitivity analysis showing the most important reactions is presented for both the benzene and toluene reacting systems. The NASA Lewis toluene mechanism's modeling capability is found to be equivalent to that of the previously published mechanism which contains a somewhat different benzene submodel.
The design optimization of air-breathing propulsion engine concepts has been accomplished by soft-coupling the NASA Engine Performance Program (NEPP) analyzer with the NASA Lewis multidisciplinary optimization tool COMETBOARDS. Engine problems, with their associated design variables and constraints, were cast as nonlinear optimization problems with thrust as the merit function. Because of the large number of mission points in the flight envelope, the diversity of constraint types, and the overall distortion of the design space; the most reliable optimization algorithm available in COMETBOARDS, when used by itself, could not produce satisfactory, feasible, optimum solutions. However, COMETBOARDS' unique features-which include a cascade strategy, variable and constraint formulations, and scaling devised especially for difficult multidisciplinary applications-successfully optimized the performance of subsonic and supersonic engine concepts. Even when started from different design points, the combined COMETBOARDS and NEPP results converged to the same global optimum solution. This reliable and robust design tool eliminates manual intervention in the design of air-breathing propulsion engines and eases the cycle analysis procedures. It is also much easier to use than other codes, which is an added benefit. This paper describes COMETBOARDS and its cascade strategy and illustrates the capabilities of the combined design tool through the optimization of a high-bypass- turbofan wave-rotor-topped subsonic engine and a mixed-flow-turbofan supersonic engine.
A general purpose optimization tool for multidisciplinary applications, which in the literature is known as COMETBOARDS, is being developed at NASA Lewis Research Center. The modular organization of COMETBOARDS includes several analyzers and state-of-the-art optimization algorithms along with their cascading strategy. The code structure allows quick integration of new analyzers and optimizers. The COMETBOARDS code reads input information from a number of data files, formulates a design as a set of multidisciplinary nonlinear programming problems, and then solves the resulting problems. COMETBOARDS can be used to solve a large problem which can be defined through multiple disciplines, each of which can be further broken down into several subproblems. Alternatively, a small portion of a large problem can be optimized in an effort to improve an existing system. Some of the other unique features of COMETBOARDS include design variable formulation, constraint formulation, subproblem coupling strategy, global scaling technique, analysis approximation, use of either sequential or parallel computational modes, and so forth. The special features and unique strengths of COMETBOARDS assist convergence and reduce the amount of CPU time used to solve the difficult optimization problems of aerospace industries. COMETBOARDS has been successfully used to solve a number of problems, including structural design of space station components, design of nozzle components of an air-breathing engine, configuration design of subsonic and supersonic aircraft, mixed flow turbofan engines, wave rotor topped engines, and so forth. This paper introduces the COMETBOARDS design tool and its versatility, which is illustrated by citing examples from structures, aircraft design, and air-breathing propulsion engine design.
The effect of the unsteady flow field in a multistage compressor on the time-averaged performance was assessed. The energy transport by the unsteady deterministic flow field was taken into account. The magnitude of the body force resulting from the aerodynamic loading on a blade row was compared to the gradient of the stress tensor associated with the unsteady time-resolved flow field generated by the blade row. The magnitude of the work performed by these forces was compared to the divergence of the energy correlations produced by the unsteady time-resolved flow field. The stress tensor and the energy correlations are non-negligible in the end wall regions. The results suggest that the turbulence is the primary source of flow mixing in the midspan region.
Currently, NASA and its industry partners are performing ground testing of hydrogen-fueled rocket based combined-cycle flowpaths. Successful ramjet and scramjet testing at Mach 6 and scramjet testing at Mach 8 have been performed. Cold flow mixing tests have also been successfully performed as have inlet operability tests. Additional testing of RBCC flowpaths at air augmented rocket and rocket only modes is underway.
In aviation gas turbine combustors, many factors, such as the degree and extent of fuel/air mixing and fuel vaporization achieved prior to combustion, influence the formation of undesirable pollutants. To assist in analyzing the extent of fuel/air mixing, flow visualization techniques have been used to interrogate the fuel distributions during subcomponent tests of lean-burning fuel injectors. Discrimination between liquid and vapor phases of the fuel was determined by comparing planar laser-induced fluorescence (PLIF) images, elastically-scattered light images, and phase/Doppler interferometer measurements. Estimates of Sauter mean diameters are made by ratioing PLIF and Mie scattered intensities for various sprays, and factors affecting the accuracy of these estimates are discussed. Mie calculations of absorption coefficients indicate that the fluorescence intensities of individual droplets are proportional to their surface areas, instead of their volumes, due to the high absorbance of the liquid fuel for the selected excitation wavelengths.
The Nozzle Acoustic Test Rig (NATR) was developed to provide additional test capabilities at Lewis needed to meet HSR program goals. The NATR is a large f ree-jet facility (free-jet diameter = 53 in.) with a design Mach number of 0.3. It is located inside a geodesic dome, adjacent to the existing Powered Lift Facility (PLF). The NATR allows nozzle concepts to be acoustically assessed for far-field (approximately 50 feet) noise characteristics under conditions simulating forward flight. An ejector concept was identified as a means of supplying the required airflow for this free-jet facility. The primary stream is supplied through a circular array of choked nozzles and the resulting low pressure in the constant, annular- area mixing section causes a "pumping" action that entrains the secondary stream. The mixed flow expands through an annular diffuser and into a plenum chamber. Once inside the plenum, the flow passes over a honeycomb/screen combination intended to remove large disturbances and provide uniform flow. The flow accelerates through an elliptical contraction section where it achieves a free-jet Mach number of up to 0.3.
An initial assessment of a proposed High-Speed Civil Transport (HSCT) was conducted in the fall of 1995 at the NASA Langley Research Center. This configuration, known as the Industry Reference-H (Ref.-H), was designed by the Boeing Aircraft Company as part of their work in the High Speed Research program. It included a conventional tail, a cranked-arrow wing, four mixed-flow turbofan engines, and capacity for transporting approximately 300 passengers. The purpose of this assessment was to evaluate and quantify operational aspects of the Reference-H configuration from a pilot's perspective with the additional goal of identifying design strengths as well as any potential configuration deficiencies. This study was aimed at evaluating the Ref.-H configuration at many points of the aircraft's envelope to determine the suitability of the vehicle to accomplish typical mission profiles as well as emergency or envelope-limit conditions. Pilot-provided Cooper-Harper ratings and comments constituted the primary vehicle evaluation metric. The analysis included simulated real-time piloted evaluations, performed in a 6 degree of freedom motion base NASA Langley Visual-Motion Simulator, combined with extensive bath analysis. The assessment was performed using the third major release of the simulation data base (known as Ref.-H cycle 2B).
A performance assessment of eight low-boom high speed civil transport (HSCT) configurations and a reference HSCT configuration has been performed. Although each of the configurations was designed with different engine concepts, for consistency, a year 2005 technology, 0.4 bypass ratio mixed-flow turbofan (MFTF) engine was used for all of the performance assessments. Therefore, all original configuration nacelles were replaced by a year 2005 MFRF nacelle design which corresponds to the engine deck utilized. The engine thrust level was optimized to minimize vehicle takeoff gross weight. To preserve the configuration's sonic-boom shaping, wing area was not optimized or altered from its original design value. Performance sizings were completed when possible for takeoff balanced field lengths of 11,000 ft and 12,000 ft, not considering FAR Part 36 Stage III noise compliance. Additionally, an arbitrary sizing with thrust-to-weight ratio equal to 0.25 was performed, enabling performance levels to be compared independent of takeoff characteristics. The low-boom configurations analyzed included designs from the Boeing Commercial Airplane Group, Douglas Aircraft Company, Ames Research Center, and Langley Research Center. This paper discusses the technology level assumptions, mission profile, analysis methodologies, and the results of the assessment. The results include maximum lift-to-drag ratios, total fuel consumption, number of passengers, optimum engine sizing plots, takeoff performance, mission block time, and takeoff gross weight for all configurations. Results from the low-boom configurations are also compared with a non-low-boom reference configuration. Configuration dependent advantages or deficiencies are discussed as warranted.
The NASA Engine Performance Program (NEPP) can configure and analyze almost any type of gas turbine engine that can be generated through the interconnection of a set of standard physical components. In addition, the code can optimize engine performance by changing adjustable variables under a set of constraints. However, for engine cycle problems at certain operating points, the NEPP code can encounter difficulties: nonconvergence in the currently implemented Powell's optimization algorithm and deficiencies in the Newton-Raphson solver during engine balancing. A project was undertaken to correct these deficiencies. Nonconvergence was avoided through a cascade optimization strategy, and deficiencies associated with engine balancing were eliminated through neural network and linear regression methods. An approximation-interspersed cascade strategy was used to optimize the engine's operation over its flight envelope. Replacement of Powell's algorithm by the cascade strategy improved the optimization segment of the NEPP code. The performance of the linear regression and neural network methods as alternative engine analyzers was found to be satisfactory. This report considers two examples-a supersonic mixed-flow turbofan engine and a subsonic waverotor-topped engine-to illustrate the results, and it discusses insights gained from the improved version of the NEPP code.
This is the final report on the research performed under NASA Glen grant NASA/NAG-3-1975 concerning feedback control of the Pratt & Whitney (PW) STF 952, a twin spool, mixed flow, after burning turbofan engine. The research focussed on the design of linear and gain-scheduled, multivariable inner-loop controllers for the PW turbofan engine using H-infinity and linear, parameter-varying (LPV) control techniques. The nonlinear turbofan engine simulation was provided by PW within the NASA Rocket Engine Transient Simulator (ROCETS) simulation software environment. ROCETS was used to generate linearized models of the turbofan engine for control design and analysis as well as the simulation environment to evaluate the performance and robustness of the controllers. Comparison between the H-infinity, and LPV controllers are made with the baseline multivariable controller and developed by Pratt & Whitney engineers included in the ROCETS simulation. Simulation results indicate that H-infinity and LPV techniques effectively achieve desired response characteristics with minimal cross coupling between commanded values and are very robust to unmodeled dynamics and sensor noise.