Hypersonic Airbreathing Research Articles

Design of hypersonic flight vehicles: some lessons from the past and future challenges

Design challenges of four basic hypersonic flight vehicle classes are discussed: non-winged (capsules) and winged re-entry vehicles, airbreathing cruise/acceleration and ascent/re-entry vehicles. Basically flight in the Earth atmosphere is considered, no systematic review is given, propulsion systems are not considered. Three topics are treated in detail: (1) The concept of the thermal state of the vehicle surface, which encompasses both thermal surface effects and thermal loads. Thermal surface effects are important especially for hypersonic airbreathing flight vehicles. (2) The aerodynamic performance (L/D) of non-winged and winged re-entry vehicles as important mission parameter. It is sketched, how the shapes of the vehicles evolved and how L/D increases were accomplished. (3) Air-vehicle engineering issues of large airbreathing hypersonic vehicles. High design sensitivities and small payload fractions of such vehicles make necessary new development and test approaches, with numerical multidisciplinary simulation and optimization as well as experimental vehicles playing a deciding role.

Numerical Simulation of Liquid Fueled SCRAMJET Combustor Flow Fields

Numerical simulations were carried out for various reacting flow fields related to scramjet propulsion system using commercial CFD Software. Three-dimensional Navier stokes equations were solved along with the K-e turbulence model. Modeling of the turbulence chemistry interaction is done through infinitely fast rate chemical kinetics. The software was validated extensively by comparing different experimental conditions for scramjet combustor to find its error band and range of application. Good agreement between the experimental and computational values were obtained for the scramjet combustor flow field with strut, pylon and cavity injection system with both hydrogen and hydrocarbon fuel. The validated CFD tool was applied in the design exercise of flight-sized kerosene fuel scramjet combustor of a hypersonic airbreathing mission. Significant improvement of combustion efficiency and thrust could be achieved by relocating the fuel injection system through the analysis of various thermochemical variables i...

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.

Preliminary Sizing of an Hypersonic Airbreathing Airliner

The purpose of this paper is to identify, for given technology levels (TRL) and mission requirements, those parameters that are critical for preliminary sizing of a hypersonic airbreathing airliner. Mission requirements will dictate a solution space of possible vehicle architecture capable of meeting cruise conditions as well as of taking-off (TO) and landing. In practice, once defined a range of cruise vehicle architectures, constraints are imposed (as to all passenger airliners), such as: 1. take off (=TO) and landing distance (so-called field length, FL): FL no longer than for the B-747-400, or 10000 ft; 2. completing TO with one engine off; 3. max acceleration at TO and climb-out (CO) = 0.4 g; 4. Hydrogen fuel (Meeting NOx emission limits (EINOx) is a further constraint not discussed in this paper). These constraints enable focusing on a realistic design out of the broad range of vehicles capable of performing the given mission. Thus a realistic vehicle must not only integrate aerodynamics and propulsion system; in fact, it is the result of many iterations in the design space, until performance and constraints are successfully achieved and met. The Gross Weight at Take Off (TOGW) was deliberately discarded as a constraint, based on Previous studies by Czysz. Typically, limiting from the beginning the TOGW leads to a vicious spiral where weight and propulsion system requirements keep growing, eventually denying convergence. In designing passenger airliners, in fact, it is the payload that is assumed fixed from the start, not the total weight. A parametric analysis of the hypersonic vehicle architecture is presented: in particular, optimal size, weight and geometrical shape are defined for different mission requirements. This analysis has shown that, it is possible to define a range of possible successful solutions for the European LAPCAT II project.

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.

Evolutionary Algorithm Shape Optimization of a Hypersonic Flight Experiment Nose Cone

Hypersonic vehicle development, particularly for hypersonic airbreathers, will require robust-design optimization to achieve performance targets. Vehicle shape optimization will play an important role. This paper presents the first application to hypersonic vehicle shape optimization of a reduced parametric section shape representation coupled with a surrogate-assisted evolutionary-algorithm optimization approach. The particular case considered is the minimization of total drag of the nose cone of a hypersonic flight experiment. The computational fluid dynamics solver used here is ANSYS CFX. Single-point optimization at Mach 3 and Mach 8 is performed at altitudes relevant to those Mach numbers for a typical hypersonic flight experiment ascent trajectory. Without surrogate assistance, the optimized shapes for each Mach number were both found to result in significant drag-force reductions (1.39% at Mach 3 and 1.96% at Mach 8) when compared with the baseline blunted standard ogive nose-cone shape. When the surrogate assistance (which is a radial-basis-function network approximation to the drag dependence on the shape parameters) was introduced, the optimized shapes yielded nearly the same drag reduction as without surrogate assistance, but also resulted in significant savings in the computational cost. Finally, the performance of the optimum shape derived at Mach 3 is evaluated at Mach 8 and vice versa to illustrate the robustness of the nose-cone shapes derived using such an approach.

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.

Control-Oriented Modeling of an Air-Breathing Hypersonic Vehicle

Full simulation models for flexible air-breathing hypersonic vehicles include intricate couplings between the engine and flight dynamics, along with complex interplay between flexible and rigid modes, resulting in intractable systems for nonlinear control design. In this paper, starting from a high-fidelity model, a control-oriented model in closed form is obtained by replacing complex force and moment functions with curve-fitted approximations, neglecting certain weak couplings, and neglecting slower portions of the system dynamics. The process itself allows an understanding of the system-theoretic properties of the model, and enables the applicability of model-based nonlinear control techniques. Although the focus of this paper is on the development of the control-oriented model, an example of control design based on approximate feedback linearization is provided. Simulation results demonstrate that this technique achieves excellent tracking performance, even in the presence of moderate parameter variations. The fidelity of the truth model is then increased by including additional flexible effects, which render the original control design ineffective. A more elaborate model with an additional actuator is then employed to enhance the control authority of the vehicle, required to compensate for the new flexible effects, and a new design is provided.

Trajectory Optimization and Conceptual Study of Small Test Vehicles for a Hypersonic Engine Using a High-Altitude Balloon

The Japan Aerospace Exploration Agency, JAXA, announced a long-term vision recently. In the vision, JAXA aims to develop hypersonic aircrafts. A pre-cooled turbojet engine has great potential as one of newly developed hypersonic airbreathing engines. We also expect the engine to be installed in space transportation vehicles in the future. For combustion test in the real flight conditions of the engines, JAXA has an experimental plan where a small test vehicle is released from a high-altitude balloon. This paper applies numerical analysis and optimization techniques to conceptual designs of the test vehicle in order to obtain the best configuration and trajectory for the flight test. The results show helpful knowledge for designing prototype vehicles.

X-43A Hypersonic vehicle technology development

NASA recently completed two major programs in hypersonics: Hyper-X, with the record-breaking flights of the X-43A, and the next generation launch technology (NGLT) program. The X-43A flights, the culmination of the Hyper-X program, were the first-ever examples of a scramjet engine propelling a hypersonic vehicle and provided unique, convincing, detailed flight data required to validate the design tools needed for design and development of future operational hypersonic airbreathing vehicles. Concurrent with Hyper-X, NASA's NGLT program focused on technologies needed for future revolutionary launch vehicles. The NGLT was “competed” by NASA in response to the President's redirection of the agency to space exploration, after making significant progress towards maturing technologies required to enable airbreathing hypersonic launch vehicles. NGLT quantified the benefits, identified technology needs, developed airframe and propulsion technology, chartered a broad University base, and developed detailed plans to mature and validate hypersonic airbreathing technology for space access. NASA is currently in the process of defining plans for a new hypersonic technology program. Details of that plan are not currently available. This paper highlights results from the successful Mach 7 and 10 flights of the X-43A, and the current state of hypersonic technology.

Investigation of Two-Stage-to-Orbit Airbreathing Launch-Vehicle Configurations

A conceptual design study was performed to assess and compare the parameters of two-stage-to-orbit airbreathing launch-vehicle systems in an effort to identify optimal configurations for improved access to space. The following vehicle configuration categories were considered: horizontal takeoff ramjet-scramjet boosters with upper-stage reusable rockets, horizontal takeoff turbine boosters with ramjet-scramjet upper stages, and vertical takeoff rocket boosters with ramjet-scramjet upper stages. Ramjet-scramjet upper stages utilize integrated rockets for final orbital ascent. Ramjet-scramjet booster stages make use of either integrated turbines or rockets for takeoff and acceleration to ramjet start. The payload requirement for each vehicle system is 20,000 lb delivered to a 100-n mile low Earth orbit. The vehicle solutions were evaluated utilizing several figures of merit including empty weight, wetted area, actively cooled area, gross weight, staging considerations, and design and technology traceability to future development of single-stage-to-orbit hypersonic airbreathing vehicles. The application of these criteria to the closed vehicle solutions reveals that the vertically launched rocket booster with an upper-stage ramjet-scramjet is the best two-stage airbreathing launch vehicle configuration. Of all of the two-stage vehicles investigated, those which place the hypersonic propulsion elements as part of the upper stage surpass those with first-stage placement and exhibit more design traceability with single-stage-to-orbit airbreathing launch vehicles.

Time evolution and mixing characteristics of hydrogen and ethylene transverse jets in supersonic crossflows

We report an experimental investigation that reveals significant differences in the near-flowfield properties of hydrogen and ethylene jets injected into a supersonic crossflow at a similar jet-to-freestream momentum flux ratio. Previously, the momentum flux ratio was found to be the main controlling parameter of the jet’s penetration. Current experiments, however, demonstrate that the transverse penetration of the ethylene jet was altered, penetrating deeper into the freestream than the hydrogen jet even for similar jet-to-freestream momentum flux ratios. Increased penetration depths of ethylene jets were attributed to the significant differences in the development of large-scale coherent structures present in the jet shear layer. In the hydrogen case, the periodically formed eddies persist long distances downstream, while for ethylene injection, these eddies lose their coherence as the jet bends downstream. The large velocity difference between the ethylene jet and the freestream induces enhanced mixing at the jet shear layer as a result of the velocity induced stretching-tilting-tearing mechanism. These new observations became possible by the realization of high velocity and high temperature freestream conditions which could not be achieved in conventional facilities as have been widely used in previous studies. The freestream flow replicates a realistic supersonic combustor environment associated with a hypersonic airbreathing engine flying at Mach 10. The temporal evolution, the penetration, and the convection characteristics of both jets were observed using a fast-framing-rate (up to 100 MHz) camera acquiring eight consecutive schlieren images, while OH planar laser-induced fluorescence was performed to verify the molecular mixing.

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.

Experimentation, Test, and Evaluation Requirements for Future Airbreathing Hypersonic Systems

Experiences gained in recent airbreathing hypersonic vehicle development programs have ascertained that mastering hypersonic e ight will depend directly on continued ability to experiment, test, and evaluate. Indeed, many of the hypersonic e ight systems being studied will require testing and modeling capabilities that do not currently exist or that may soon be lost if current infrastructures are not maintained. The testing and evaluation of hypersonic systems represents a unique set of challenges. Duplication of the hypersonic e ight environment requires extreme temperatures and pressures coupled with complex physical interactions. Ground-test facilities are limited in their ability to duplicate all salient parameters simultaneously. Data from e ight experimentation are also limited due to range airspace requirements for long distance e ight corridors. These limitations manifest themselvesin rangesafety, data acquisition, and environmental concerns.Computational techniquesarea growing supplement to experiment, but are not a replacement. Despite the improvement in both hardware and algorithms, analytical models and computational techniques are extremely time consuming, fall short of adequate e delity, and requiredata to anchor andvalidatethem.Theconclusion isthat only through a uniquecombination of experiment, e ight testing, analysis, and computation, will hypersonic system risks be reduced and the vision of airbreathing hypersonic e ight be realized.

Thermodynamic Cycle Analysis of Magnetohydrodynamic-Bypass Hypersonic Airbreathing Engines

The prospects for realizing a magnetohydrodynamic (MHD) bypass hypersonic airbreathing engine are examined from the standpoint of fundamental thermodynamic feasibility. The MHD-bypass engine, first proposed as part of the Russian AJAX vehicle concept, is based on the idea of redistributing energy between various stages of the propulsion system flow train. The system uses an MHD generator to extract a portion of the aerodynamic heating energy from the inlet and an MHD accelerator to reintroduce this power as kinetic energy in the exhaust stream. In this way, the combustor entrance Mach number can be limited to a specified value even as the flight Mach number increases. Thus, the fuel and air can be efficiently mixed and burned within a practical combustor length, and the flight Mach number operating envelope can be extended. In this paper, we quantitatively assess the performance potential and scientific feasibility of MHD-bypass engines using a simplified thermodynamic analysis. This cycle analysis, based on a thermally and calorically perfect gas, incorporates a coupled MHD generator-accelerator system and accounts for aerodynamic losses and thermodynamic process efficiencies in the various engin components. It is found that the flight Mach number range can be significantly extended; however, overall performance is hampered by non-isentropic losses in the MHD devices.