Results for “RAMJET ENGINE” — Page 9

NASA NTRS1994

Shock-tunnel combustor testing for hypersonic vehicles

Proposed configurations for the next generation of transatmospheric vehicles will rely on air breathing propulsion systems during all or part of their mission. At flight Mach numbers greater than about 7 these engines will operate in the supersonic combustion ramjet mode (scramjet). Ground testing of these engine concepts above Mach 8 requires high pressure, high enthalpy facilities such as shock tunnels and expansion tubes. These impulse, or short duration facilities have test times on the order of a millisecond, requiring high speed instrumentation and data systems. One such facility ideally suited for scramjet testing is the NASA-Ames 16-Inch shock tunnel, which over the last two years has completed a series of tests for the NASP (National Aero-Space Plane) program at simulated flight Mach numbers ranging from 12-16. The focus of the experimental programs consisted of a series of classified tests involving a near-full scale hydrogen fueled scramjet combustor model in the semi-free jet method of engine testing whereby the compressed forebody flow ahead of the cowl inlet is reproduced (see appendix A). The AIMHYE-1 (Ames Integrated Modular Hypersonic Engine) test entry for the NASP program was completed in April 1993, while AIMHYE-2 was completed in May 1994. The test entries were regarded as successful, resulting in some of the first data of its kind on the performance of a near full scale scramjet engine at Mach 12-16. The data was distributed to NASP team members for use in design system verification and development. Due to the classified nature of the hardware and data, the data reports resulting from this work are classified and have been published as part of the NASP literature. However, an unclassified AIAA paper resulted from the work and has been included as appendix A. It contains an overview of the test program and a description of some of the important issues.

Loomis, Mark P.↗

NASA NTRS2000

THRSTER: A Three-Stream Ejector Ramjet Analysis and Design Tool

An engineering tool for analyzing ejectors in rocket based combined cycle (RBCC) engines has been developed. A key technology for multi-cycle RBCC propulsion systems is the ejector which functions as the compression stage of the ejector ramjet cycle. The THRee STream Ejector Ramjet analysis tool was developed to analyze the complex aerothermodynamic and combustion processes that occur in this device. The formulated model consists of three quasi-one-dimensional streams, one each for the ejector primary flow, the secondary flow, and the mixed region. The model space marches through the mixer, combustor, and nozzle to evaluate the solution along the engine. In its present form, the model is intended for an analysis mode in which the diffusion rates of the primary and secondary into the mixed stream are stipulated. The model offers the ability to analyze the highly two-dimensional ejector flowfield while still benefits from the simplicity and speed of an engineering tool. To validate the developed code, wall static pressure measurements from the Penn-State and NASA-ART RBCC experiments were used to compare with the results generated by the code. The calculated solutions were generally found to have satisfactory agreement with the pressure measurements along the engines, although further modeling effort may be required when a strong shock train is formed at the rocket exhaust. The range of parameters in which the code would generate valid results are presented and discussed.

Chue, R. S.↗

NASA NTRS2000

THRSTER: A THRee-STream Ejector Ramjet Analysis and Design Tool

An engineering tool for analyzing ejectors in rocket based combined cycle (RBCC) engines has been developed. A key technology for multi-cycle RBCC propulsion systems is the ejector which functions as the compression stage of the ejector ramjet cycle. The THRee STream Ejector Ramjet analysis tool was developed to analyze the complex aerothermodynamic and combustion processes that occur in this device. The formulated model consists of three quasi-one-dimensional streams, one each for the ejector primary flow, the secondary flow, and the mixed region. The model space marches through the mixer, combustor, and nozzle to evaluate the solution along the engine. In its present form, the model is intended for an analysis mode in which the diffusion rates of the primary and secondary into the mixed stream are stipulated. The model offers the ability to analyze the highly two-dimensional ejector flowfield while still benefits from the simplicity and speed of an engineering tool. To validate the developed code, wall static pressure measurements from the Penn-State and NASA-ART RBCC experiments were used to compare with the results generated by the code. The calculated solutions were generally found to have satisfactory agreement with the pressure measurements along the engines, although further modeling effort may be required when a strong shock train is formed at the rocket exhaust. The range of parameters in which the code would generate valid results are presented and discussed.

Chue, R. S.↗

NASA NTRS1996

A CFD Study of Turbojet and Single-Throat Ramjet Ejector Interaction

Supersonic ejector-diffuse systems have application in driving an advanced airbreathing propulsion system, consisting of turbojet engines acting as the primary and a single throat ramjet acting as the secondary. The turbojet engines are integrated into the single throat ramjet to minimize variable geometry and eliminate redundant propulsion components. The result is a simple, lightweight system that is operable from takeoff to high Mach numbers. At this high Mach number (approximately Mach 3.0), the turbojets are turned off and the high speed ramjet/scramjet take over and drive the vehicle to Mach 6.0. The turbojet-ejector-ramjet system consists of nonafterburning turbojet engines with ducting canted at 20 degrees to supply supersonic flow (downstream of CD nozzle) to the horizontal ramjet duct at a supply total pressure and temperature. Two conditions were modelled by a 2-D full Navier Stokes code at Mach 2.0. The code modelled the Fabri choke as well as the non-Fabri non critical case, using a computational throat to supply the back pressure. The results, which primarily predict the secondary mass flow rate and the mixed conditions at the ejector exit were in reasonable agreement with the 1-D cycle code (TBCC).

Chang, Ing↗

NASA NTRS1986

Orbit-on-demand vehicle propelled by air-turborocket/ramjet engines

A preliminary design study has been completed for a fully reusable, single-stage-to-orbit transatmospheric vehicle. The specified mission capability was to lift a 20,000 lb payload to low earth orbit. A ground accelerator-assisted horizontal take-off was chosen to increase operational flexibility. The multi-mode propulsion system included the use of air-turborocket, ramjet, scramjet and rocket engines. Weight and performance estimates were obtained for the vehicle. A computer package was developed to perform aerothermodynamic analyses of the propulsion modes throughout the flight environment from take-off to low earth orbit. Results are presented for a semi-optimized trajectory. The analysis indicates that a vehicle of this type has great potential for providing low cost, flexible access to space.

Hartung, L.↗

NASA NTRS2010

Enriched Storable Oxidizers for Rocket Engines

The name "enriched storable oxidizers" (ESOs) has been coined for a family of optimized mixtures of between two and four oxidizer fluids. For most applications, the constituents of these mixtures would be nitrogen tetroxide (N2O4), nitrous oxide (N2O), and nitrogen dioxide (NO2); in some applications, the mixtures might include inhibited red fuming nitric acid [IRFN (which consists of red fuming nitric acid to which some hydrogen fluoride is added to reduce its corrosive effect]. The optimum proportions of these constituents would be different for different applications. ESOs were originally proposed for use in spacecraft and launch-rocket propulsion systems: ESOs could be especially useful in advanced spacecraft propulsion systems that could operate in multiple modes. ESOs might also be useful in special terrestrial applications that could include ramjet and scramjet aircraft engines.

Sackheim, R. L.↗

NASA NTRS2002

Hypersonic Propulsion at Pratt and Whitney: Overview

Pratt & Whitney (P&W) is developing the technology for hypersonic components and engines. A supersonic combustion ramjet (scramjet) database was developed during the National Aero Space Plane (NASP) program using hydrogen fueled propulsion systems for space access vehicles and serves as a point of departure for the current emphasis on hydrocarbon scramjets. The Air Force Hypersonic Technology (HyTech) Office has put programs in place to develop the technologies necessary to demonstrate the operability, performance and structural durability of a liquid hydrocarbon fueled scramjet system that operates from Mach 4 to 8. Fuel-cooled superalloys and lightweight structures are being developed to improve thermal protection and durability and to reduce propulsion system weight. The application of scramjet engine technology as part of combined cycle propulsion systems is also being pursued under NASA and U.S. Air Force sponsorship. The combination of scramjet power and solid rocket booster acceleration is applicable to hypersonic cruise missiles. Scramjets that use gas turbines for low speed acceleration and scramjets using rocket power for low speed acceleration are being studied for application to reusable launch systems and hypersonic cruise vehicles. P&W's recent activities and future plans for hypersonic propulsion will be described.

Kazmar, Richard R.↗

NASA NTRS1989

Optimal guidance and propulsion control for transatmospheric vehicles

Problems associated with on-board trajectory optimization and with the synthesis of guidance laws are addressed for ascent to LEO of an air-breathing, single-stage-to-orbit vehicle. A multimode propulsion system is assumed which incorporates turbojet, ramjet, scramjet, and rocket engines. An energy-state approximation is applied to a four-state dynamic model for flight of a point mass over a spherical nonrotating earth. An algorithm for generating fuel-optimal climb profiles is derived via singular perturbation theory. This algorithm results from application of the minimum principle to a low-order dynamic model that includes general functional dependence on angle of attack and a normal component of thrust. Switching conditions are derived which, under appropriate assumptions, govern optimal transition from one propulsion mode to another. The use of bank angle to modulate the magnitude of the vertical component of lift is shown to improve the index performance. Numerical results illustrate the nature of the resulting fuel-optimal climb paths.

Corban, J. E.↗

NASA NTRS1989

A real-time guidance algorithm for aerospace plane optimal ascent to low earth orbit

Problems of onboard trajectory optimization and synthesis of suitable guidance laws for ascent to low Earth orbit of an air-breathing, single-stage-to-orbit vehicle are addressed. A multimode propulsion system is assumed which incorporates turbojet, ramjet, Scramjet, and rocket engines. An algorithm for generating fuel-optimal climb profiles is presented. This algorithm results from the application of the minimum principle to a low-order dynamic model that includes angle-of-attack effects and the normal component of thrust. Maximum dynamic pressure and maximum aerodynamic heating rate constraints are considered. Switching conditions are derived which, under appropriate assumptions, govern optimal transition from one propulsion mode to another. A nonlinear transformation technique is employed to derived a feedback controller for tracking the computed trajectory. Numerical results illustrate the nature of the resulting fuel-optimal climb paths.

Calise, A. J.↗

NASA NTRS1990

Rapid near-optimal trajectory generation and guidance law development for single-stage-to-orbit airbreathing vehicles

General problems associated with on-board trajectory optimization, propulsion system cycle selection, and with the synthesis of guidance laws were addressed for an ascent to low-earth-orbit of an air-breathing single-stage-to-orbit vehicle. The NASA Generic Hypersonic Aerodynamic Model Example and the Langley Accelerator aerodynamic sets were acquired and implemented. Work related to the development of purely analytic aerodynamic models was also performed at a low level. A generic model of a multi-mode propulsion system was developed that includes turbojet, ramjet, scramjet, and rocket engine cycles. Provisions were made in the dynamic model for a component of thrust normal to the flight path. Computational results, which characterize the nonlinear sensitivity of scramjet performance to changes in vehicle angle of attack, were obtained and incorporated into the engine model. Additional trajectory constraints were introduced: maximum dynamic pressure; maximum aerodynamic heating rate per unit area; angle of attack and lift limits; and limits on acceleration both along and normal to the flight path. The remainder of the effort focused on required modifications to a previously derived algorithm when the model complexity cited above was added. In particular, analytic switching conditions were derived which, under appropriate assumptions, govern optimal transition from one propulsion mode to another for two cases: the case in which engine cycle operations can overlap, and the case in which engine cycle operations are mutually exclusive. The resulting guidance algorithm was implemented in software and exercised extensively. It was found that the approximations associated with the assumed time scale separation employed in this work are reasonable except over the Mach range from roughly 5 to 8. This phenomenon is due to the very large thrust capability of scramjets in this Mach regime when sized to meet the requirement for ascent to orbit. By accounting for flight path angle and flight path angle rate in construction of the flight path over this Mach range, the resulting algorithm provides the means for rapid near-optimal trajectory generation and propulsion cycle selection over the entire Mach range from take-off to orbit.

Calise, A. J.↗

NASA NTRS1991

Rapid near-optimal aerospace plane trajectory generation and guidance

Problems associated with onboard trajectory optimization, propulsion system cycle selection, and the synthesis of guidance laws are addressed for ascent to low earth orbit of an airbreathing, single-stage-to-orbit vehicle. A multicycle propulsion system is assumed that incorporates turbojet, ramjet, scramjet, and rocket engines. An energy state approximation is applied to a singularly perturbed, four-state dynamic model for flight of a point mass over a spherical nonrotating earth. An algorithm is then derived for generating both the fuel-optimal climb profile and the guidance commands required to follow that profile. In particular, analytic switching conditions are derived that, under appropriate assumptions, efficiently govern optimal transition from one propulsion cycle to another. The algorithm proves to be computationally efficient and suitable for real-time implementation. The paper concludes with the presentation of representative numerical results that illustrate the nature of the fuel-optimal climb paths and the tracking performance of the guidance algorithm.

Corban, J. E.↗

NASA NTRS1992

A numerical study of mixing in supersonic combustors with hypermixing injectors

A numerical study was conducted to evaluate the performance of wall mounted fuel-injectors designed for potential Supersonic Combustion Ramjet (SCRAM-jet) engine applications. The focus of this investigation was to numerically simulate existing combustor designs for the purpose of validating the numerical technique and the physical models developed. Three different injector designs of varying complexity were studied to fully understand the computational implications involved in accurate predictions. A dual transverse injection system and two streamwise injector designs were studied. The streamwise injectors were designed with swept ramps to enhance fuel-air mixing and combustion characteristics at supersonic speeds without the large flow blockage and drag contribution of the transverse injection system. For this study, the Mass-Averaged Navier-Stokes equations and the chemical species continuity equations were solved. The computations were performed using a finite-volume implicit numerical technique and multiple block structured grid system. The interfaces of the multiple block structured grid systems were numerically resolved using the flux-conservative technique. Detailed comparisons between the computations and existing experimental data are presented. These comparisons show that numerical predictions are in agreement with the experimental data. These comparisons also show that a number of turbulence model improvements are needed for accurate combustor flowfield predictions.

Lee, J.↗

NASA NTRS1993

A numerical study of mixing in supersonic combustors with hypermixing injectors

A numerical study was conducted to evaluate the performance of wall mounted fuel-injectors designed for potential Supersonic Combustion Ramjet (SCRAM-jet) engine applications. The focus of this investigation was to numerically simulate existing combustor designs for the purpose of validating the numerical technique and the physical models developed. Three different injector designs of varying complexity were studied to fully understand the computational implications involved in accurate predictions. A dual transverse injection system and two streamwise injector designs were studied. The streamwise injectors were designed with swept ramps to enhance fuel-air mixing and combustion characteristics at supersonic speeds without the large flow blockage and drag contribution of the transverse injection system. For this study, the Mass-Average Navier-Stokes equations and the chemical species continuity equations were solved. The computations were performed using a finite-volume implicit numerical technique and multiple block structured grid system. The interfaces of the multiple block structured grid systems were numerically resolved using the flux-conservative technique. Detailed comparisons between the computations and existing experimental data are presented. These comparisons show that numerical predictions are in agreement with the experimental data. These comparisons also show that a number of turbulence model improvements are needed for accurate combustor flowfield predictions.

Lee, J.↗

NASA NTRS1993

Mathematical description of complex chemical kinetics and application to CFD modeling codes

A major effort in combustion research at the present time is devoted to the theoretical modeling of practical combustion systems. These include turbojet and ramjet air-breathing engines as well as ground-based gas-turbine power generating systems. The ability to use computational modeling extensively in designing these products not only saves time and money, but also helps designers meet the quite rigorous environmental standards that have been imposed on all combustion devices. The goal is to combine the very complex solution of the Navier-Stokes flow equations with realistic turbulence and heat-release models into a single computer code. Such a computational fluid-dynamic (CFD) code simulates the coupling of fluid mechanics with the chemistry of combustion to describe the practical devices. This paper will focus on the task of developing a simplified chemical model which can predict realistic heat-release rates as well as species composition profiles, and is also computationally rapid. We first discuss the mathematical techniques used to describe a complex, multistep fuel oxidation chemical reaction and develop a detailed mechanism for the process. We then show how this mechanism may be reduced and simplified to give an approximate model which adequately predicts heat release rates and a limited number of species composition profiles, but is computationally much faster than the original one. Only such a model can be incorporated into a CFD code without adding significantly to long computation times. Finally, we present some of the recent advances in the development of these simplified chemical mechanisms.

Bittker, D. A.↗

NASA NTRS1993

Mathematical Description of Complex Chemical Kinetics and Application to CFD Modeling Codes

A major effort in combustion research at the present time is devoted to the theoretical modeling of practical combustion systems. These include turbojet and ramjet air-breathing engines as well as ground-based gas-turbine power generating systems. The ability to use computational modeling extensively in designing these products not only saves time and money, but also helps designers meet the quite rigorous environmental standards that have been imposed on all combustion devices. The goal is to combine the very complex solution of the Navier-Stokes flow equations with realistic turbulence and heat-release models into a single computer code. Such a computational fluid-dynamic (CFD) code simulates the coupling of fluid mechanics with the chemistry of combustion to describe the practical devices. This paper will focus on the task of developing a simplified chemical model which can predict realistic heat-release rates as well as species composition profiles, and is also computationally rapid. We first discuss the mathematical techniques used to describe a complex, multistep fuel oxidation chemical reaction and develop a detailed mechanism for the process. We then show how this mechanism may be reduced and simplified to give an approximate model which adequately predicts heat release rates and a limited number of species composition profiles, but is computationally much faster than the original one. Only such a model can be incorporated into a CFD code without adding significantly to long computation times. Finally, we present some of the recent advances in the development of these simplified chemical mechanisms.

Bittker, D. A.↗

NASA NTRS2004

Exo-Skeletal Engine: Novel Engine Concept

The exo-skeletal engine concept represents a new radical engine technology with the potential to substantially revolutionize engine design. It is an all-composite drum-rotor engine in which conventionally heavy shafts and discs are eliminated and are replaced by rotating casings that support the blades in spanwise compression. Thus the rotating blades are in compression rather than tension. The resulting open channel at the engine centerline has immense potential for jet noise reduction and can also accommodate an inner combined-cycle thruster such as a ramjet. The exo-skeletal engine is described in some detail with respect to geometry, components, and potential benefits. Initial evaluations and results for drum rotors, bearings, and weights are summarized. Component configuration, assembly plan, and potential fabrication processes are also identified. A finite element model of the assembled engine and its major components is described. Preliminary results obtained thus far show at least a 30-percent reduction of engine weight and about a 10-dB noise reduction, compared with a baseline conventional high-bypass-ratio engine. Potential benefits in all aspects of this engine technology are identified and tabulated. Quantitative assessments of potential benefits are in progress.

Chamis, Cristos C.↗

NASA NTRS1974

Abating exhaust noises in jet engines

A noise abating improvement for jet engines including turbojets, turbofans, turboprops, ramjets, scramjets, and hybrid jets is introduced. A provision is made for an apparatus in the primary and/or secondary flow streams of the engines; the apparatus imparts to the exhaust gases a component rotation or swirl about the engine's longitudinal axis. The rotary component in the exhaust gases causes a substantial suppression of sound energy build up normally produced by an axial flow exhaust system.

Schwartz, I. R.↗

NASA NTRS1997

Future Aerospace Technology in the Service of the Alliance: Sustained Hypersonic Flight - Volume 3

Sustained Hypersonic Flight capability is envisioned for several future military and civilian applications, such as long range immediate reaction reconnaissance, high speed interception of air targets, long range precision strike against hardened or time critical targets, and access to space. The inherent reduction in time-to-target and low vulnerability will permit new operational tactics. The symposium outlined mission needs and operational scenarios for hypersonic vehicles. Addressed were: Technological issues and challenges in external hypersonic aerodynamics and design, propulsion and engine/airframe integration, military ramjet applications, overall system design including structures and materials development needs, and test facilities. Ground based test facilities cannot provide full simulation at Mach numbers above 5. Therefore, the use of Computational Fluid Dynamics design tools is essential. For routine use of the computational tools in analysis, design and optimization, it is necessary to reduce the time for the entire computational process by two orders of magnitude. Hypersonic boundary layer transition remains a critical design issue because of the important impact on engine drag and on heating, which can affect the choice of materials and thermal protection systems. In addition, reliable prediction of scramjet net thrust is an absolute must in resolving hypersonic air breathing vehicle design issues. Due to current facility and computational shortfalls, the development of future hypersonic flight systems requires research flight tests in the technology areas of boundary layer transition and air-breathing propulsion engine performance. For sustained hypersonic flight beyond Mach 6, the supersonic combustion ramjet (scramjet) engine is the only choice for the near future. Only this air-breathing concept offers a significant promise of large reductions in required propellant fractions, increased payload fractions, and reduced size vehicles, together with a foreseeable technological feasibility. Airframe/engine integration, combustor design and thermal management are the predominant engineering tasks. Fuels, hydrogen or hydrocarbon, must be matched to the operational needs of military or civil use. Experience in existing ramjet propelled missiles capable of speeds up to Mach 4 can support the development effort. The potential mission and cost benefits of sustained hypersonic flight to both military and civil applications are tremendous. From the budget point of view, the possibility of sharing development costs between military and civil programs offers a specific advantage.

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