Hypersonic Airbreathing Propulsion Research Articles

1708 軸対称型スクラムジェットエンジンのインレット始動方法の数値的研究(OS17-2 超音速流と衝撃波,オーガナイズドセッション)

Hypersonic airbreathing propulsion, in particular, scramjets (supersonic combustion ramjets), is a promising technology that offers great potential for efficient and economical access-to-space and atmospheric transport. Reliable in-flight inlet starting is of critical importance for the successful operation of scramjet engines, particularly internal-compression configurations with high-contraction inlets. A wide variety of methods including diaphragm rupture and sliding doors (diaphragm erosion) have been contrived in order to address the inlet starting issue, which is inherent to hypersonic airbreathing systems with internal compression. This paper discusses the mechanism and causes of inlet unstart from aerodynamic viewpoints and assesses the effectiveness and viability of various starting methods, based on steady and transient simulations by means of high-fidelity computational fluid dynamics.

Orbital Delivery of Small Payloads Using Hypersonic Airbreathing Propulsion

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Numerical simulation of air/H<SUB align=right>2 combustion processes in a scramjet turbulent flow

The aim of this work is to develop a numerical approach, based on the Favre's averaged Navier-Stokes equations coupled with a turbulence model and with a complex finite rate Air/Hydrogen combustion kinetic model to simulate the unsteady axisymetric supersonic hydrogen-air mixing processes. The main application is dedicated to the scramjet hypersonic air-breathing propulsion engine. To achieve this goal the Spalart-Allmaras turbulence model (SA) with correction terms to model the compressibility effects and the complete finite rate chemistry model of Jachimowsky involving 13 species and 153 reactions have been implemented in a finite volume code.

A new concept and preliminary study of variable hypersonic inlet with fixed geometry based on shockwave control

The inlet plays a significant role in the hypersonic airbreathing propulsion. A fixed-geometry inlet is troubled by low air-capture ratio and large additional drag while operating below the design Mach number. Whereas a variable-geometry inlet can maximize performance, but adds weight and complexity to the propulsion system. Based on a fluidic shock shape control technique, this paper proposes a new concept of variable hypersonic inlet with fixed geometry, gives the realization scheme, and conducts a preliminary validation. The results show that the control of the external shock system and the effective throat area can be achieved by the self-provided high pressure fluid of the inlet. For an inlet with an operating Mach-number region of 4 to 6, the shock-on-lip condition can be maintained from Mach 5 to Mach 6 with the maximum expense of 1.8% secondary flow ratio, resulting in 20% extra captured mass flow and 8% less forebody drag at low Mach numbers compared with conventional fixed-geometry inlets. Thus, the performance enhancement by using the proposed variable inlet can substantially benefit the acceleration process of hypersonic vehicles at low Mach numbers.

Continue development of air breathing hypersonic propulsion

Continue development of air breathing hypersonic propulsion

Preliminary X-43 flight test results

The successful Mach 7 flight test of the Hyper-X/X-43 research vehicle has provided a major, essential demonstration of the capability of the airframe integrated scramjet engine. This flight was a crucial first step toward establishing the potential for air-breathing hypersonic propulsion for application to space-launch vehicles.

Hyper-X: Foundation for future hypersonic launch vehicles

The successful Mach-7 flight test of the Hyper-X/X-43A research vehicle has provided a major, essential demonstration of the capability of the airframe-integrated scramjet engine and hypersonic airbreathing vehicle design tools and vision vehicles. This flight was a crucial step toward establishing air-breathing hypersonic propulsion for application to space-launch vehicles and other hypersonic systems. This paper examines the significance of the flight test in advancing the state-of-the science and provides a strategic vision for achieving the dream for safe, efficient and reliable space access with air-breathing propulsion in the near future, through use of more near term approaches.

NASA hypersonic flight demonstrators—overview, status, and future plans

NASA's Next Generation Launch Technology (NGLT) program is developing and maturing advanced propulsion and vehicle systems technologies and flight vehicle concepts to enable future development of safer and more economical launch systems. Within NGLT, NASA is developing advanced air breathing propulsion systems and demonstrating these systems in hypersonic flight vehicles. The flight demonstrations are necessary to fully validate these technologies for application to future space launch vehicles and other flight systems. NASA's Hyper-X Program (X-43A) began the effort to flight demonstrate hypersonic air breathing propulsion systems to provide technologies that will enable development of safer and more economic space access vehicles in the future. Following X-43A, NASA, in collaboration with the United States (US) Department of Defense (DoD), is developing additional, progressively more complex hypersonic X-vehicles that will demonstrate new air breathing propulsion systems, propulsion-airframe integration, and other vehicle systems technologies required for high speed flight up to Mach 15. These technologies will contribute to safer, more reliable and more economic future launch systems and hypersonic aircraft/missiles. This paper describes NASA's current hypersonic flight demonstration projects, status of the efforts, and plans for future vehicles.

Validation of the Wilcox k-omega Model for Flows Characteristic to Hypersonic Airbreathing Propulsion

Numerical results obtained with the WARP code solving the 1988 Wilcox k‐ω two-equation model are compared with experimental data of flowfields characteristic of hypersonic airbreathing propulsion. The problems chosen for the comparison include a shock/boundary-layer interaction case, a reacting and inert planar mixing case, and two ramp injector mixing cases. In addition, a comparison is performed with empirical correlations on the basis of skin friction for flow over a flat plate and shear layer growth for a free shear layer. The agreement between the numerical and experimental results varies between being reasonable and excellent, with a discrepancy generally not exceeding 20%. It is found that the grid-induced error can be reduced to an acceptable level for most problems with a reasonable mesh size. However, the free shear layer and the shock/boundary-layer interaction cases require a considerably finer mesh. The Wilcox dilatational dissipation correction is seen to be beneficial in predicting the growth of a free shear layer at a high convective Mach number, but its use is considered either questionable or detrimental for the other cases. A proper choice of the turbulent Schmidt number is observed to be crucial in predicting the injectant mole fraction contours for one of the ramp injector cases, with the best agreement obtained with Schmidt number Sct fixed to 0.25. For the inert planar mixing case, overall better agreement is obtained when setting both the turbulent Prandtl and Schmidt number to 0.5.

초음속 풍동의 추력 측정 방법 비교

극초음속 추진기관의 설계 및 평가에 있어 추력의 측정은 매우 중요하다. 일반적인 추력 측정방법으로는 추력 측정기를 사용하고 있지만 이러한 방법으론 엔진의 자유제트 실험이나 모델 연소기 같은 실험에는 적용이 부적절하다. 이 때문에 피토압력을 이용한 새로운 추력 계산 방법을 고려하였고 검증하였다. 피토 압력을 통해 계산된 추력의 검증은 추력측정기를 통해 측정된 실제 추력값과의 비료를 통해 이루어질 수 있으며 이를 위해 추력측정기 위에 소형 초음속 풍동을 장착하였다. 추력이 측정되는 동안 충동 후방의 피토압력을 동시에 측정하였다. 측정된 피토 압력을 이용하여 구한 추력값을 검증하기 위해서, 추력측정기로 측정한 추력을 완정팽창 노즐이론 및 추력계수를 이용한 이론적인 계산에 의해 구해진 추력값과 비료, 분석하였다. The determination of thrust is essential in design and evaluation of a hypersonic airbreathing propulsion device. Conventional methods to determine the thrust is using thrust stand or force measurement system. However, these conventional methos are not applicable to the case where thrusts stands are impractical, such as free jet testing of engines, and model combustor. For this reason, the thrust determination method from measured pitot pressure is considered and validated. Validation of thrust determination from pitot pressures can be achieved by comparing the actual thrust from thrust stand. For validation purpose, a small-scale supersonic wind tunnel is installed on the thrust stand. Thrusts are measured while pressures are measured simulaneously. Then, the thrust from pitot pressure measurements are compared with the measured thrust and theoretical thrusts.

NASA's Integrated Space Transportation Plan — 3 rd generation reusable launch vehicle technology update

NASA's Integrated Space Transportation Plan (ISTP) calls for investments in Space Shuttle safety upgrades, second generation Reusable Launch Vehicle (RLV) advanced development and third generation RLV and in-space research and technology. NASA's third generation launch systems are to be fully reusable and operation by 2025. The goals for third generation launch systems are to reduce cost by a factor of 100 and improve safety by a factor of 10,000 over current systems. The Advanced Space Transportation Program Office (ASTP) at NASA's Marshall Space Flight Center in Huntsville, AL has the agency lead to develop third generation space transportation technologies. The Hypersonics Investment Area, part of ASTP, is developing the third generation launch vehicle technologies in two main areas, propulsion and airframes. The program's major investment is in hypersonic airbreathing propulsion since it offers the greatest potential for meeting the third generation launch vehicles. The program will mature the technologies in three key propulsion areas, scramjets, rocket-based combined cycle and turbine-based combination cycle. Ground and flight propulsion tests are being planned for the propulsion technologies. Airframe technologies will be matured primarily through ground testing. This paper describes NASA's activities in hypersonics. Current programs, accomplishments, future plans and technologies that are being pursued by the Hypersonics Investment Area under the Advanced Space Transportation Program Office will be discussed.

Characterization of Expansion Tube Flows for Hypervelocity Combustion Studies

An expansion tube with an 89 mm inner diameter and a 12 m length was characterized for basic studies of the near-e eld mixing and e ame-holding characteristics of gaseous fuels injected in high-stagnation-enthalpy supersonic e ows. The expansion tube provides a wide range of freestream conditions with little perturbation of test gas chemical composition. The latter is critical for supersonic-combustion studies in the high-stagnationenthalpy e ows associated with hypersonic airbreathing propulsion systems. Efforts focused on achieving three operatingpoints,simulatinge ightMach8,10,and12stagnation-enthalpyconditionsattheentranceofasupersonic combustor. The ability of the expansion tube to provide a steady-e ow test time of adequate duration and a core e ow of sufe cient size for 2-mm jet-in-crosse ow studies was verie ed. A core-e ow size of 25 mm could be achieved in 6.2 MJ/kg stagnation-enthalpy e ows with 400 πs of steady test time (e ight Mach 12 simulation ). The primary effects of the boundary layer on the expansion tube e ow were to extend the useful test time (in contrast with conventional shock-tube performance ) and to decrease the core-e ow size.

Numerical Simulation of Hypersonic Shock-Induced Combustion Ramjets

Hypersonic air-breathing propulsion utilizing shock-induced combustion ramjets is investigated. Twodimensional geometries are simulated with planar and axisymmetric configurations, as well as external and mixed-compression configurations. The lower-upper Symmetric Gauss-Seidel scheme combined with a symmetric shock-capturing total variation diminishing scheme are used to solve the Euler equations, with nonequilibrium chemical reactions. The finite rate chemistry model includes 13 species (H2, O2, H, O, OH, H2O, HO2, H2O2, N, NO, HNO, N2, and NO2) and 33 reactions. The numerical method has been verified by comparison with H2/air induction delay times, analytical solutions to wedge problems, and exothermic blunt body flows. Results obtained with an inviscid, chemically nonequilibrium numerical approach and with realistic geometries demonstrate that shock-induced combustion can be used as a viable means of hypersonic propulsion. As part of the combustor design, it has also been numerically demonstrated that a minimum-entrop y, or Chapman-Jouguet condition exists for oblique-detonation waves generated by wedges in nonequilibrium chemically reacting H2/air flowfields.

Errata: Strategy for Developing Air-Breathing Aerospace Planes

Acknowledgment T paper is an updated and shortened version of AirBreathing Aerospace Plane Development Essential: Hypersonic Propulsion Flight Tests (Invited)/' published in the Proceedings of the 2nd European Symposium on Aerothermodynamics for Space Vehicles, held at ESTEC, Noordwijk, Netherlands on Nov. 21-23, 1994 (ESA SP-367, Feb. 1995). References 4 and 21 should read as follows: Heiser, W. H., and Pratt, D. T., with Daly, D. H., and Mehta, U. B., Hypersonic Airbreathing Propulsion, AIAA Education Series, AIAA, Washington, DC, 1994. Augenstein, B. W., and Harris, E. D., with Aroesty, J., Blumenthal, I., Brooks, N., Frelinger, D., Garber, T., Gottenmoeller, R., Hiland, J., Liu, S. K., Pace, S., Rosen, J., Rowell, L., and Stucker, J., The National Aerospace Plane (NASP): Development Issues for the Follow-on Vehicle, Executive Summary, RAND, R3878/1-AF, 1993.

Analysis of concepts for single stage to orbit

Introduction T HE Space Shuttle was designed in the 1970s and is still the most efficient manned Earth-to-orbit vehicle available. However, operation of the Space Shuttle costs much more than planned. For future space missions, a new transportation system is desired to improve the operability and costs. The primary purpose of this study was to consider configurations for a new vehicle. Many ideas have been introduced. For example, the U.S. Air Force's concept Black Horse is a single-stage horizontal-takeoff spacecraft with aerial refueling, and Pegasus is a four-stage Earth-to-orbit vehicle using a reusable airplane as the first stage. Another example is the National Aerospace Plane concept, which used hypersonic airbreathing propulsion. This Note presents the results for two concepts designed to place small payloads in orbit with fully reusable vehicles. For these concepts, a vertical takeoff and horizontal landing (VTHL) single-stage-to-orbit (SSTO) concept and an assisted SSTO using a subsonic airplane for air launch are compared.

Hypersonic Airbreathing Propulsion. W.H. Heiser and D.T. Pratt with D.H. Daley and U.B. Mehta. American Institute of Aeronautics and Aeronautics, 370 L'Enfant Promenade SW, Washington, DC 20024-2518, USA. 1994. 587 pp plus diskette. Illustrated. $69.96 (AIAA members), $89.95 (non-members).

Hypersonic Airbreathing Propulsion. W.H. Heiser and D.T. Pratt with D.H. Daley and U.B. Mehta. American Institute of Aeronautics and Aeronautics, 370 L'Enfant Promenade SW, Washington, DC 20024-2518, USA. 1994. 587 pp plus diskette. Illustrated. 89.95 (non-members). - Volume 99 Issue 986

Perspective on hypersonic nonequilibrium flow

Introduction O as well as new hypersonic flow programs continue to challenge the aerospace engineer. These programs are often of consuming difficulty, involving the synthesis of chemical kinetics, quantum mechanics, and radiation physics with fluid dynamics. To further complicate matters, the flowfield is often rarefied; thus the Knudsen number requires independent consideration along with the Mach and Reynolds numbers. Although these parameters are related, different regions of the flow, such as the interior of a bow shock, depend differently on them as compared with other regions. Because of its complexity, recent research has focused on the modeling of more realistic, complex hypersonic flows with the intensive use of computational fluid dynamics (CFD). Of necessity, the large codes used in this effort must resort to various empiricisms and approximations. Our objective in this Survey is to discuss some of these empiricisms and approximations within the context of hypersonic nonequilibrium flow. The issues and advances examined are primarily chosen for their relevance to theoretical gasdynamics; however, this Review makes no claim to being comprehensive. The books by Anderson and Park provide background material for the discussion of current issues, including lessons learned from the design and operation of the Space Shuttle. Our presentation is partly based on a previous report, some of which was published elsewhere. An alternative discussion of some of the same material can be found in an article by Tirsky whose discussion is especially valuable for emphasizing hypersonic research by Russian workers, with greater emphasis given to engineering heat transfer predictions. Some of this work is not as well known in the western community as it should be. The importance of finite rate chemistry and the problem of fuel-air mixing to hypersonic airbreathing propulsion cannot be overemphasized. These topics, however, are omitted from our discussion in as much as they fall under combustion gasdynamics and ought to be examined in this context.' The treatise on hydrogen combustion presented in Ref. 8 should be of interest to scramjet propulsion researchers. A broad perspective on hypersonic airbreathing vehicle design may be found in a survey (to be three volumes) edited by Bertin et al. In the next section, nonequilibrium aerothermodynamics is discussed. Much of this section is concerned with energy exchange that involves the vibrational modes of diatomic molecules. The subsequent section reviews the modeling of rarefied hypersonic flows and their continuum extension. There is no shortage of approaches that are available, as well as open issues. These approaches encompass the direct simulation Monte Carlo (DSMC) method, NavierStokes (NS) equations, Burnett, or augmented Burnett equations, and the 13-moment equations. The review concludes with a few brief remarks.

Facility opportunities and associated stream chemistry considerations for hypersonic air-breathing propulsion

Wallace Chinitz is a Principal Scientist at General Applied Science Laboratories (GASL) as well as a Professor of Mechanical Engineering at Cooper Union's Nerken School of Engineering. He received his Ph.D. from the Polytechnic Institute of Brooklyn in 1962, and taught at that institution before leaving to conduct research at GASL until 1972. Dr. Chinitz began his career at the Fairchild Engine Division in 1957, then worked at Republic Aviation Corporation until 1960. He has published over 100 articles, papers, and reports. Dr. Chinitz is an Associate Fellow of the AIAA and is a member of the ASME, the ASEE, and the Combustion Institute.

Wave combustors for trans-atmospheric vehicles

A performance analysis is given of a conceptual transatmospheric vehicle (TAV). The TAV is powered by a an oblique detonation wave engine (ODWE). The ODWE is an airbreathing hypersonic propulsion system which utilizes shock and detonation waves to enhance fuel-air mixing and combustion in supersonic flow. In this wave combustor concept, an oblique shock wave in the combustor can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture, thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. The result is a shorter lighter engine compared to the scramjet. The ODWE-powered hypersonic vehicle performance is compared to that of a scramjet-powered vehicle. Among the results outlined, it is found that the ODWE trades a better engine performance above Mach 15 for a lower performance below Mach 15. The overall higher performance of the ODWE results in a 51,000-lb weight savings and a higher payload weight fraction of approximately 12 percent.

Hypersonic turbomachinery-based air-breathing engines for the earth-to-orbit vehicle

Hypersonic air-breathing engines will make the Earth-to-orbit vehicle completely different from the present one powered by rocket engines. The space plane propelled by a certain hypersonic air-breathing propulsion system is expected to appear in the next century. The turbomachinery-based engine (turboengine) is a candidate for the space plane propulsion system and will be combined with scramjet and rocket engines. Turboengines, including turboramjet, air-turboramjet, and their modifications, may be applied as the accelerators to the space plane having a high specific impulse at a rather low supersonic Mach number. Here, a conceptual study of these turboengines with preliminary system design, performance calculations, and consideration of relative merits of the engine concepts is performed for the configuration, performance, weight, and size. An engine evaluation with mission capability of the space plane for assumed requirements is made. As a result, engine performance depends on the liquid oxygen utilization, and weight and size of the engine are important factors for application to the space plane. Thus a certain optimization of the engine system itself and of a combination of the engines would be necessary.