Scramjet Technology Research Articles

Optimization and data mining for shock-induced mixing enhancement inside scramjet using stochastic deep-learning flowfield prediction

One of the significant challenges in supersonic combustion ramjet engines lies in effective mixing of the fuel, primarily due to the high momentum of supersonic inflow at the combustor. Among the various fluid phenomena associated with fuel mixing, it is well recognized that the interaction between the fuel-injection jet plume and oblique shock waves can significantly enhance mixing efficiency. However, the most suitable interaction for optimal mixing enhancement remains yet to be clarified. The present study conducts model-based optimization and data mining for shock-induced mixing enhancement of angled-slot injection in a two-dimensional scramjet combustor. Stochastic deep-learning flowfield prediction has been utilized to enable fast and reliable evaluations of a substantial number of designs. Prediction errors can be estimated without requiring correct data owing to uncertainty quantification techniques. Data mining and sensitivity analysis, coupled with flowfield prediction, have revealed the optimal shock interaction with the fuel jet plume characterized by a pronounced downstream recirculation region. The mechanism that drives mixing enhancement through this recirculation region has been discussed based on the results of optimization and sensitivity analysis. This study has yielded valuable insights for the future design of scramjet injectors. Furthermore, the effectiveness of the model-based design and analysis has been demonstrated through the present study, showcasing its potential for guiding future developments in scramjet technology.

Correction to: Aero-structural Analysis of a Scramjet Technology Demonstrator Designed to Operate at an Altitude of 23 km at Mach 5.8

Correction to: Aero-structural Analysis of a Scramjet Technology Demonstrator Designed to Operate at an Altitude of 23 km at Mach 5.8

Aero-structural Analysis of a Scramjet Technology Demonstrator Designed to Operate at an Altitude of 23 km at Mach 5.8

Aero-structural Analysis of a Scramjet Technology Demonstrator Designed to Operate at an Altitude of 23 km at Mach 5.8

Multi-objective optimization of a hypersonic airbreathing vehicle

Multi-objective optimization of a hypersonic airbreathing engine (scramjet technology) was carried out with the aim of maximizing thrust and minimizing drag while satisfying a series of design constraints, such as avoiding unstart (blockage of supersonic flow within the combustion chamber) by ensuring that the pressure ratio across the shock waves remains below the adverse pressure gradient given by the Korkegi limit, geometry correction to achieve shock on-lip condition, and temperature and pressure requirements at the inlet exit. Using the relations presented in the literature, pressure and viscous drag are estimated analytically. The analytical approach is verified against computational fluid dynamics data from Ansys Fluent to solve two-dimensional compressible Reynolds-averaged Navier–Stokes flow equations, with transition shear stress transport as the turbulence closure model. Comparing the total drag and the flow properties at the combustion chamber entrance shows the model's feasibility for the optimization approach. Three different approaches were conducted to formulate the multi-objective function to determine the one that can find the highest number of geometries satisfying the Korkegi limit with the highest net thrust. The best approach was the multi-objective function formulated with the uninstalled thrust, total pressure recovery, and pressure drag, concentrating the search in the region with greater uninstalled thrust and lower drag and nearly doubling the value of net thrust compared to the first formulation, which uses the uninstalled thrust, pressure drag, and viscous drag.

Impact of passive fuel injections techniques in the flow field of the scramjet combustor

• The introduction of passive injection techniques in a parallel cavities scramjet combustor influences the internal flow filed charateristics. • The effect of fuel injection location on the internal flow field charaterictics is also investigated. • Out of all the configuration , the scrmajet combustor showef optimum performance when the fuel is injected through the wall only. The development of scramjet technology has paved the path of possibilities for researchers to develop light-weighted space missions in the years to come. Many countries around the globe have already tested successful scramjet engines. The researchers are making continuous efforts to improve the combustor design by improvising turbulence inserts and flame holders in the combustor. One such improvisation is the introduction of strut into the parallel-cavities combustor. The research paper uses computational fluid dynamics to investigate the effect of internal flow field variables on a strut and cavity-based scramjet combustor. The current work presents a comparative study by analyzing two different scramjet combustor configurations namely case A and Case B to understand the effect of presence of strut in the cavity combustor. The velocity and static pressure profiles are analyzed at different streamwise locations. At x = 70 mm, the velocity values are reduced to lower magnitude for case B as compared to case A and the static pressure values towards downstream detoriates with reduction in intensity of shock waves. It is seen that the magnitude of velocity values and pressure values increases with increase is Mach number. The high-pressure values are observed for M-2.75. A performance assessment of the combustor is carried out for both the configurations, maximum combustion and mixing performance is obtained for M-2.52 for case B as compared to Case A. The paper further analyses three different fuel injection locations namely fuel injection strut only, fuel injection through the wall only and fuel injection through both strut+wall recognized as the case I, II, and III respectively. From various flow-field parameters analysis, it is seen that the production of eddies and vortices in the combustor region indicates proper mixing which is observed for case I. On studying the mass fraction concentration of H 2 0 species for all the cases it is observed that the H 2 0 species is increasing in the lateral direction which indicates that the combustion efficiency is incrementing in the same direction. The maximum combustion and mixing efficiency is observed for case I than the other two cases.

Influence of wall mounted ramps on DLR strut scramjet combustor under non-reacting flow field

The scramjet technology is considered to be the advanced engine for high-speed transportations. The present numerical examination addresses the influence of double ramp's implications in the scramjet engine combustor walls accompanied by strut injection technique under non-reacting flow conditions. The numerical analyses are performed using simulation software, ANSYS Fluent 18.0. Computational problems have been modeled using Reynolds averaged Navier-Stokes (RANS) equations. The inlet flow condition of the model is Mach (Ma) 2 and sonic speed hydrogen fuel from the strut is driven along the flow direction. The characteristics of flow for the two-dimensional (2D) combustor models are studied using shadowgraph images and Mach number contours. Numerical solutions are analogized with the published experimental outcome for validation. The wall static pressure values are higher at the vicinity of the strut ramps regions due to the additional shock interactions generated from the ramps compared to the DLR scramjet model. The subsonic zones are created due to additional shock waves and enhance air–fuel mixing for ramp-mounted combustors. The total pressure loss across the combustor increases as the ramps are positioned downstream of the strut.

Preliminary experimental study on solid-fuel rocket scramjet combustor

Liquid or gaseous fuel scramjet technology has made great progress, and there has been some research attention to solid-fuel scramjet. A new scramjet configuration using solid fuel as propellant, namely solid-fuel rocket scramjet, is tested experimentally. It consists of two combustors. One is a rocket combustor used as gas generator, and the other is a supersonic combustor used for secondary combustion. The experiment simulates a flight Mach number of 4 at high altitude (stagnation temperature and pressure are 1170 K and 1.16 MPa, respectively), and metalized solid fuel is used as propellant. The results reveal that fuel-rich gas from the gas generator can burn with air in the supersonic combustor. Preliminary evaluation results show that the combustion efficiency of the propellant is about 90%, and the total pressure recovery coefficient in the supersonic combustor is about 0.6. These results indicate that the configuration of solid-fuel rocket scramjet is feasible.

Ground-based real-time auto commanding system for ISRO advanced technology vehicles

In general, space transportation systems are expendable, resulting in huge launch cost. To meet the future challenge of low-cost access to space, the Indian Space Research Organisation (ISRO) envisaged plans to develop systems that are recoverable and reusable, besides adopting effcient propulsion systems like air-breathing rockets. Advanced Technology Vehicle (ATV) D01 is ISRO's new-generation high-performance sounding rocket vehicle offering a cost-effective test bed for demonstrating air-breathing propulsion. Experiments with scramjet technology testing are carried out at Satish Dhawan Space Centre (SDSC) SHAR, Sriharikota. For this experiment, ATV-D01 is confgured as a two-stage, un-guided, spinning, fn-stabilized vehicle with two protruding passive scramjet engines.

Design of an Airbreathing Second Stage for a Rocket-Scramjet-Rocket Launch Vehicle

The most promising alternative to rockets for improved access to space involves staged systems using airbreathing propulsion. With scramjet technology improving, a number of airbreathing assisted access-to-space vehicle concepts have recently been proposed, including a three-stage rocket-scramjet-rocket launch architecture for payload masses on the order of 100 kg. This article presents a design methodology developed for the airbreathing second stage of such a system. This methodology uses multidisciplinary design optimization with simplified methods for the calculation of vehicle aerodynamics, propulsion, and mass. It has been applied to the design of a reusable scramjet-powered winged cone vehicle with a near-term Mach 6–12 hydrogen-fueled scramjet for propulsion. Through the manipulation of five vehicle design parameters, including the size and position of the engines, and flying the vehicle along constant dynamic pressure trajectories, a configuration was developed to maximize payload mass fraction to low Earth orbit. The mass of the whole system was 11.3 t, and that of the airbreathing vehicle was 4625 kg, delivering a 211 kg payload to low Earth orbit. This corresponds to an overall payload mass fraction of 1.87%. This methodology supplies a useful framework for developing a better understanding of the key drivers for airbreathing hypersonic accelerators.

Supersonic liquid fuel jets injected into quiescent air

Supersonic fuel jets may have applications in diesel engine and scramjet technology. Their properties require a fundamental examination as their mixing characteristics are likely to be affected by the leading edge shock wave. Such jets have been created experimentally in the laboratory but require further CFD studies to examine details that are obscure in the experiments. This paper reports on the CFD methods used. These include a solid body assessment of the formation of the leading edge shock, the steady state and transient analysis of the mixing layer in supersonic vapour and liquid jets. A liquid jet, transient solution is computationally expensive even when atomization and vaporization are not included. Some typical results are outlined.

Thermal Analysis of Fluid-Structural Interaction in High-Speed Engine Flowfields

The impetus toward development of a hydrogen-fueled scramjet engine to accelerate aerospace vehicles at hypersonic speeds has focused attention on the need to model accurately the fluid-thermal-structural interaction of such engines. The proposed method is able to solve coupled thermal problems by computing high-speed turbulent and compressible flowfields including multidimensional heat conduction in adjacent walls. To achieve a csompletety conservative coupling at the fluid-structural interface, the same finite element method is applied to both the fluid an the structural equations. The use of the surface energy-balance equation with the boundary integral formulation yields a stable solution procedure for steady-state computations of the stiff problem. For the computation of the heat transfer inside a supersonic combustion chamber, the nonreactive gas flow solver is extended by an energy source term to simulate the heat addition due to the combustion process. The code is tested and applied to a combustor and test conditions experimentally investigated within the framework of the joint German-Russian cooperation on scramjet technology.

Strutjet Engine Performance

To increase the technology readiness level of hypersonic scramjet technology, an innovative strut-based dualmoderamjetenginedesign capableofpoweringa hypersonicvehicleatspeedsfrom Mach4to 8isinvestigated.Two versions of the engine are under development: a dual-mode (ramjet/scramjet )engine formissile applicationsand a rocket-based combined cycle engine for space access. This paper will discuss the features of the dual-mode engine design and how they contribute to overall engine performance. Engine component performance was evaluated using a combination of analysis and component testing. Inlet performance data (mass capture, pressure recovery, and inlet/isolator pressure ratio ) were obtained from wind-tunnel testing and correlated with model and full-scale computational e uid dynamics (CFD)analyses. Combustorperformance data (combustion efe ciency, fuel-injection performance, and pressure distributions ) were obtained from full-scale direct-connect combustor testing. Nozzle losses and efe ciency were obtained from CFD analysis. Performance parameters derived from these component tests and analyses were fed into a standard hypersonic engine cycle code to predict e ight engine performance. The demonstrated component performance values were extrapolated to determine predicted component performance at the end of engine development, which were used to compute the vision e ight engine performance.

Air-Breathing Hypersonic Cruise: Prospects for Mach 4–7 Waverider Aircraft

There is currently a renewal of world-wide interest in hypersonic flight. Vehicle concepts being considered range from cruise missiles to SSTO and TSTO vehicles. The new characteristics of these vehicles are that they will be powered by air-breathing engines and have long residence times in the air-breathing corridor. In the Mach 4–7 regime, waverider aircraft are being considered as candidates for both long-range and short-range cruise missions, as hypersonic missiles, and as high-L/D highly maneuverable vehicles. This paper will discuss the potential for near-term and far-term application of air-breathing engines to the above-mentioned waverider vehicle concepts and missions. In particular, the cruise mission is discussed in detail and attempts are made to compare and contrast it with the accelerator mission. Past criticisms levied against waveriders alleging low volumetric efficiency, lack of engine/airframe integration studies, poor off-design performance, poor take-off and landing capability, have been shown by ongoing research to be unfounded. A discussion is presented of some of the technical challenges and ongoing research aimed at realizing such vehicles: from turboramjet and scramjet technology development, propulsion-airframe integration effects on vehicle performance, aeroservothermoelastic systems analysis, hypersonic stability and control with aeroservothermoelastic and propulsion effects, etc. A unique and very strong aspect of hypersonic vehicle design is the integration and interaction of the propulsion system, aerodynamics, aerodynamic heating, stability and control, and materials and structures. This first-order multidisciplinary situation demands the ability to integrate highly coupled and interacting elements in a fundamental and optimal fashion to achieve the desired performance. Some crucial technology needs are found in propulsion-airframe integration and its role in configuration definition, hypersonic boundary-layer transition and its impact on vehicle gross-weight and mission success, scramjet combustor mixing length and its impact on engine weight and, CFD (turbulence modeling, transition modeling, etc) as a principal tool for the design of hypersonic vehicles. Key technology implications in thermal management, structures, materials, and flight control systems will also be briefly discussed. It is concluded that most of the technology requirements in the Mach 4–7 regime are relatively conventional, making cited applications near-term, yet offering very significant advancements in aircraft technology.

Integrated Scramjet Nozzle/Afterbody Performance Analysis

A two-dimensiona l flowfield computer program developed for preliminary performance analysis of integrated scramjets is presented. Scramjet flowfields are computed by means of the shock-capture technique, with real gas thermodynamic properties for frozen and equilibrium compositions. The program computes internal and external flowfields with multiple shock interactions. Interaction of the exhaust (underexpanded or overexpanded) with the external stream and vehicle afterbody also is considered. Forces and moments are computed due to stream thrust and external surface pressure distributions. Typical program results for parametric nozzle/afterbody design and off-nominal performance analysis are presented. N recent years, increased attention has been given to the application of supersonic combustion ramjets (scramjets) to airbreathing launch vehicles, advanced civil aviation, and hypersonic research aircraft. One of the scramjet technology areas requiring particular attention is the interaction of the airbreathing propulsion system with the vehicle aerodynamics and the resulting effects on the overall vehicle performance. Particularly significant and as yet not well defined is the interaction of the vehicle afterbody and the exhaust of the integrated scramjet.M The integrated scramjet propulsion system utilizes the vehicle afterbody as an extension of the nozzle, producing increased effective exhaust velocities and thrust at the expense of incurred trim penalties. Therefore, a requirement exists for the means to define accurately scramjet nozzle/afterbody exhaust flowfields. A critical review of this problem indicated that detailed analysis should include consideration of real gas three-dimensional flowfields with chemical-kinetic effects, multiple shocks, heat transfer, boundary layer growth, and interaction with the inviscid flow. In view of the complexity of the problem, simplified practical methods of analysis must be developed which retain the most essential features of the flow phenomena, fulfill preliminary design requirements, and produce realistic results. At the present time, the computer programs of Refs. 5 and 6 appear to be the only published programs specifically developed to meet this requirement. These programs consider inviscid and adiabatic quasi-three-dimensional flowfields with real gas equilibrium hydrogen-air chemistry. These programs are based on the reference plane technique, with a table lookup thermodynamic properties routine. This paper represents a simpler scramjet flowfield analysis computer program suitable for preliminary design analyses. Since the majority of the current scramjet configurations are of the modular quasi-two-dimensional type, it is felt that for preliminary design .purposes the flowfield can be described sufficiently well by means of a two-dimensional program. The two-dimensional scramjet flowfield program7 considers inviscid and adiabatic supersonic flow in ducts and coflowing dissimilar streams with real or ideal gas properties. The methodology used in this program is the shock capture technique (SCT) described in Refs. 8 and 9. Thermodynamic properties are evaluated by means of a special-purpose sub