Scramjet Inlet Research Articles

Numerical Research on the NS-SDBD Control of a Hypersonic Inlet in Off-Design Mode

The overall performance of a scramjet inlet will decline while entering off-design mode. Active flow control using nanosecond surface dielectric barrier discharge (NS-SDBD) can be a novel solution to such inlet–unstart problems. NS-SDBD actuators are deployed on the surface of the internal compression section, controlling the shock waves and the separation area. Numerical simulations of hypersonic flows are carried out using the compressible Reynolds average Navier–Stokes equation (RANS), along with the plasma phenomenological model which is added in as the energy source term. Flow structures and the evolution of performance parameters are analyzed. Results show that NS-SDBD actuators are able to increase the static pressure behind the cowl shock, boosting the downstream total pressure. The compression effect becomes stronger while raising the frequency or shortening the spacing between the actuators. Under the inlet–unstart conditions, the compression wave generated by the actuator pushes the reattachment point forward, making the separation bubble longer in length and shorter in height, which reduces the strength of the separation shock. The results provide a numerical basis for the state control of a hypersonic inlet.

Influence of high temperature non-equilibrium effects on Mach 12 scramjet inlet

To better understand the high temperature non-equilibrium effects of a high Mach number (above Mach 8) scramjet inlet, the thermal and chemical non-equilibrium flow of Mach 12 two-dimensional inlet is numerically analyzed by using the thermochemical non-equilibrium gas model including a two-temperature model and air chemical reactions. The thermal equilibrium flow is simulated by using the models of thermally perfect gas and chemical non-equilibrium gas. It is found that the thermal non-equilibrium effects are relatively strong near the cowl shock and in the outer layer of the boundary layer, and gradually weaken in the downstream zone of the cowl shock. The chemical non-equilibrium effects and thermal equilibrium flows mainly exist in the high-temperature region of the boundary layer. Compared with the chemical non-equilibrium gas, the oxygen dissociation reaction near the lower wall for the thermochemical non-equilibrium gas is weaker in the external compression section; It is stronger due to fully excited vibration energy, in the internal compression section. Compared with the other models, the inlet compression ability for the thermochemical non-equilibrium gas is higher, the coefficient of mass flow and total pressure recovery for it are lower. Hence, the high temperature non-equilibrium effects can not be ignored in the design of the high Mach number scramjet inlet.

Experimental Characterization of Three-Dimensional Scramjet Inlet with Variable Internal Contraction

A compact three-dimensional scramjet intake with a high compression ratio and a movable cowl as a starting device and for the adaption of the internal contraction to respective operating conditions was designed for a parabolic reentry flight experiment. In an extensive measurement campaign, conducted in a blowdown wind tunnel at Mach 7, the operational behavior of this inlet was characterized. The parameters investigated include internal contraction ratio, Reynolds number, angle of attack, and angle of yaw, and the starting behavior was examined. Static and total pressure ratios and mass capture ratios were used to evaluate the performance of the inlet and throttle curves to determine the operational limits. The flowfield was studied by schlieren images, wall pressure distributions, and Mach number profiles in the interface of the isolator and combustion chamber, which were determined with a pressure measurement rake with static and pitot pressure probes. The flow on the external ramp was analyzed by infrared thermography. The inlet showed good performance regarding compression and efficiency and proved to be quite robust at various flight angles. The performance tended to improve, and the inlet operated more stably with increasing internal contraction, with the most important factor being whether the inlet leading-edge shock was captured by the cowl.

Three-dimensional CFD investigation of a scramjet inlet under different freestream conditions

Scramjet (supersonic combustion ramjet) engines are often seen as a promising alternative to place payloads in the Earth's orbit. Such air-breathing engines have a simple structure and few moving parts. On the other hand, high heat fluxes and pressure loads on the walls, shock-wave-boundary-layer interactions, and the risk of choked flow inside the isolator channel are a few examples of obstacles that need to be addressed during scramjet design. Therefore, this study aims to evaluate of different air freestream conditions on the flow field in a scramjet inlet (compression ramps and isolator) through detailed three-dimensional computational fluid dynamics (CFD) simulations. The Mach number, flight altitude, and angle of attack were the evaluated conditions. Moreover, this study also focuses on the behavior of the boundary layer separation located at the compression ramp corners, isolator entrance, and scramjet sidewalls. Regarding the Mach number variation, the results showed that high Mach numbers yielded high-pressure levels throughout the engine. Furthermore, a higher altitude promoted a lower total pressure field. Considering the angle of attack changes, it is evident that a higher angle of attack results in a decrease in airflow pressure, while an increase in the total pressure along the walls is observed. By investigating these freestream parameters, this work can contribute to the early phase of engine design, avoiding critical failures in the scramjet structure due to aerodynamic load and thermal stress.

Optimization of scramjet inlet based on temperature and Mach number of supersonic combustion

A two-dimensional mixed compression scramjet inlet design is presented in this work. The design is based on the static temperature and Mach number at the combustion chamber inlet according to the conditions required to burn hydrogen spontaneously at supersonic speed. This method considers air as a calorically perfect gas, with no viscous effects, and shock on-lip and shock on-corner. To burn hydrogen spontaneously at supersonic speed the mixture temperature of the income airflow and hydrogen is obtained considering the zeroth and first Laws of Thermodynamics. The criterion of equal shock strength, based on the normal component of the airflow velocity approaching the incident oblique shock waves, is applied to obtain the compression ramps angles and the airflow corresponding thermodynamic properties. Therefore, the total pressure ratio across all incident oblique shock waves and total temperature at the compression section are constants. Although the present method can be applied to any scramjet with a mixed compression system for any number of ramps, cases with up to 5 ramps, flying at a hypersonic speed corresponding to Mach number 7 through the Earth's atmosphere at 30 km of geometric altitude are considered. Finally, the same optimization criteria are extended to analyze the inlet of a scramjet vehicle with 5 compression ramps, flying at speeds from Mach numbers 5 to 10, at an altitude of 30 km.

Estimation of Aerodynamic Heating on Scramjet Inlets and Validation With Measurements

Abstract Aerodynamic heating levels on a typical inlet configuration of a scramjet engine are estimated using both standard design correlations and numerical flow simulations. The stagnation point heat flux is estimated using the Fay and Riddell formula. Aerodynamic heating over the inclined ramps is estimated using Van Driest method. Numerical flow simulations are carried out using a Reynolds averaged Navier–Stokes (RANS) solver coupled with energy equations and the Shear Stress Transport (SST) k–ω turbulence model. The aerodynamic heat flux estimates are validated with in-house measurements in a shock tunnel and for a scramjet flight experiment in the Mach number range 1.59 to 7.92. The emergence of a good agreement between them confirms the appropriateness of design correlations for heat flux estimation in scramjet inlets. The choice of simplification and appropriateness of design correlations to complex geometries demand critical assessment. Numerical flow simulations capture flow features and enable the identification of potential augmented heating zones, which will be critical for long-duration scramjet missions.

One-equation turbulence models applied to practical scramjet inlet

Abstract Reynolds-averaged Navier–Stokes equations are used to simulate a practical scramjet inlet geometry using the shock-unsteadiness modified Spalart–Allmaras (SA) turbulence model. The geometry consists of fore-body ramps, expansion corners, and inlet ducts. The focus is to study the impingement of the cowl shock on the opposite wall boundary-layer. The resulting separation bubble can lead to blockage and inlet unstarts. The shock-unsteadiness correction is employed and is found to improve the computational fluid dynamics (CFD) prediction of flow separation in shock/boundary-layer interactions. The shock-unsteadiness parameter is calibrated against available experimental data of canonical flows, and the predicted flow-field is analyzed in detail. A large separation bubble size normalized to the upstream boundary-layer thickness of 4.6 is observed in the interaction region. Across the reattachment region in the interaction region, a peak value of wall pressure is observed. The inlet performance parameters are also calculated. The total pressure losses of 62% are observed across different shock waves, with an additional loss of 15% due to viscous boundary-layer effects.

One-equation turbulence models applied to practical scramjet inlet

Abstract Reynolds-averaged Navier–Stokes equations are used to simulate a practical scramjet inlet geometry using the shock-unsteadiness modified Spalart–Allmaras (SA) turbulence model. The geometry consists of fore-body ramps, expansion corners, and inlet ducts. The focus is to study the impingement of the cowl shock on the opposite wall boundary-layer. The resulting separation bubble can lead to blockage and inlet unstarts. The shock-unsteadiness correction is employed and is found to improve the computational fluid dynamics (CFD) prediction of flow separation in shock/boundary-layer interactions. The shock-unsteadiness parameter is calibrated against available experimental data of canonical flows, and the predicted flow-field is analyzed in detail. A large separation bubble size normalized to the upstream boundary-layer thickness of 4.6 is observed in the interaction region. Across the reattachment region in the interaction region, a peak value of wall pressure is observed. The inlet performance parameters are also calculated. The total pressure losses of 62% are observed across different shock waves, with an additional loss of 15% due to viscous boundary-layer effects.

넓은 비행영역을 고려한 2D 스크램제트 흡입구 설계 방법

넓은 비행영역을 고려한 2D 스크램제트 흡입구 설계 방법

Aerodynamic Study of the NASA’s X-43A Hypersonic Aircraft

A 2D aerodynamic study of the NASA’s X-43A hypersonic aircraft is developed using two different approaches. The first one is analytical and based on the resolution of the oblique shock wave and Prandtl–Meyer expansion wave theories supported by an in-house program and considering a simplified aircraft’s design. The second approach involves the use of a Computational Fluid Dynamics (CFD) package, OpenFOAM and the real shape of the aircraft. The aerodynamic characteristics defined as the lift and drag coefficients, the aerodynamic efficiency and the pitching moment coefficient are calculated for different angles of attack. Evaluations are made for an incident Mach number of 7 and an altitude of 30 km. For both methodologies, the required angles of attack to achieve a Vertical Force Balance (VFB) and a completely zero pitching moment conditions are considered. In addition, an analysis to optimise the nose configuration of the aircraft is performed. The mass flow rate throughout the scramjet as a function of the angle of attack is also presented in the CFD model in addition to the pressure, density, temperature and Mach fields. Before presenting the corresponding results, a comparison between the aerodynamic coefficients in terms of the angle of attack of both models is carried out in order to properly validate the CFD model. The paper clarifies the requirements needed to make sure that both oblique shock waves originating from the leading edge meet just at the scramjet inlet clarifying the advantages of fulfilling such condition.

Finite-time nonsingular terminal sliding mode control-based fuzzy smooth-switching coordinate strategy for AHV-VGI

This paper presents a fuzzy smooth-switching coordinate strategy for air-breathing hypersonic vehicle with variable geometry scramjet inlet (AHV-VGI), which can improve the aerodynamic performance of hypersonic vehicle by moving the movable lip to capture more airflow. The longitudinal model of AHV-VGI is firstly formulated, followed by a strict feedback form model. Wherein, the multiple nonlinear aerodynamic models are established with different elongation distance of translating cowl. Secondly, a finite-time nonsingular terminal sliding mode control (NTSMC) method is introduced to develop velocity and altitude subsystem controllers for each model. After that, an interval type-2 fuzzy logic system technique is employed to guarantee the smoothness during switching process between the adjacent NTSMC controllers. Furthermore, the Lyapunov function combined with adaptive laws are constructed to prove the finite-time stability of the closed-looped switching system. Finally, simulation results are provided to validate the effectiveness of these theoretical results.

Comment on “A new approach for the design of hypersonic scramjet inlets” [Phys. Fluids 24, 086103 (2012)

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Response to “Comment on ‘A new approach for the design of hypersonic scramjet inlets’” [Phys. Fluids 32, 079101 (2020)

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Experimental and numerical heat transfer from vortex-injection interaction in scramjet flowfields

ABSTRACTAir-breathing propulsion has the potential to decrease the cost per kilogram for access-to-space, while increasing the flexibility of available low earth orbits. However, to meet the performance requirements, fuel-air mixing inside of scramjet engines and thermal management still need to be improved.An option to address these issues is to use intrinsically generated vortices from scramjet inlets to enhance fuel-air mixing further downstream, leading to shorter, less internal drag generating, and thus more efficient engines. Previous works have studied this vortex-injection interaction numerically, but validation was impractical due to lack of published experimental data. This paper extends upon these previous works by providing experimental data for a canonical geometry, obtained in the T4 Stalker Tube at Mach 8 flight conditions, and assesses the accuracy of numerical methodologies such as RANS CFD to predict the vortex-injection interaction.Focus is placed on understanding the ability of the numerical methodology to replicate the most important aspects of the vortex-injection interaction. Results show overall good agreement between the numerical and experimental results, as all major features are captured. However, limitations are encountered, especially due to a localised region of over predicted heat flux.

Numerical evaluation of effects of the isolator height and the leading-edge bluntness on a scramjet inlet

Air-breathing hypersonic vehicles can be powered by scramjet (supersonic combustion ramjet) engines, whose inlet requires blunt leading edges in order to overcome the high heat fluxes that are inherent of hypersonic flights. Thus, this work has the purpose of evaluating, through detailed two-dimensional computational fluid dynamics analysis, the influence of the leading-edge bluntness and isolator height on the airflow of a scramjet intake. Three radii have been analyzed for the forebody and cowl leading edges: sharp, 0.5 and 1.0 mm. Moreover, two different isolator heights were varied: 15 and 20 mm. Fixed temperatures of 300 K and 1300 K were considered along the hypersonic vehicle wall. The static temperature, pressure, and Mach number contour are part of the analysis, as well as the Stanton number and the adiabatic kinetic efficiency of the scramjet. It was evidenced that both leading-edge bluntness and isolator height performed great influence on the intake airflow structure. Results showed that the isolator height had a major effect on the flow separation at the isolator entrance yielding a recirculation zone that increases with the isolator height. Furthermore, higher Stanton numbers and temperature along the isolator were found for the shorter isolator height due to stronger shock waves and higher frequency of shock-wave train. The leading-edge bluntness, in this analysis, was taken as a parameter to reduce the heat flux at the stagnation point of leading edges. However, a decrease in the adiabatic kinetic efficiency was observed when the leading-edge radius is increased. Moreover, blunt leading edges promoted a decrease in Stanton number along the scramjet intake walls and also a decrease in average temperatures throughout the isolator.

2D Hypersonic Scramjet Inlet Geometry of Mach 7 using CFD

In the era of space transportations there is a huge demand on space technology to improve on cost reduction and take the heavy loads into space. Thus the load carrying capacities will be increase with this air breathing engines. This work gives a report on the design, analysis of optimal 2D scramjet engine inlet operating at Mach 7 without use of movable geometry. A computational study for scramjet inlet with different ramp angles is carried out. Several cases are considered to compress the air by rounding leading edge without moving the whole cowl up and down, by fixing the cowl lip and assuming axisymmetric inlet with rounded edge. The numerical tests are conducted to obtain maximum total pressure recovery, drag force and outlet Mach number for given flight condition. Two dimensional effects are studied with Navier-strokes approach to compute the pressure and Mach number at a different location. Oblique shock waves, expansion waves and shock wave interactions are primarily focused. Computational Fluid Dynamics (CFD) solver is used; steady flow simulations are carried out for inlet geometries with one, two, three and four ramps. Geometrical shape is redesigned based on oblique shock wave analysis. The corrected model is tested on Fluent with boundary layer considerations that the theoretical analysis is not able to cover. Lastly, a conclusion summarizing the design process is drawn and the optimal model is recommended for the Mach 7 inlet with different ramps with contraction ratio 10. It had been observed that two ramp scramjet inlet model has been preferred to use which has optimum pressure recovery and lower drag.

Investigation on 2D Hypersonic Scramjet Inlet of Mach 7 using CFD

Investigation on 2D Hypersonic Scramjet Inlet of Mach 7 using CFD

Effect of scramjet inlet vortices on fuel plume elongation and mixing rate

ABSTRACTHypersonic air-breathing propulsion can improve cost and flexibility of Low Earth Orbit (LEO) satellite launch missions. However, at the high flight Mach numbers required for access-to-space, performance margins are extremely tight. Techniques to improve mixing efficiency can push this technology forward. However, these are required to produce a minimal increase in losses and heat loads to be viable. The use of inlet-generated vortices in scramjets for mixing enhancement was previously studied. These vortices interact with the injected fuel plume, stretching it and increasing its effective surface for mixing. Moreover, these vortices are intrinsic to the flowfield. Therefore, contrary to other methods, when using inlet vortices mixing is enhanced without producing additional heat loads or losses. This work studies the vortex-injection interaction through numerical RANS simulations. A non-dimensional variable defining the quality of the plume shape for mixing purposes is proposed. This parameter is used to assess the effect of vortex intensity and injector location on fuel plume shape. The results show the ability of inlet vortices to modify fuel plume shape significantly increasing fuel mixing rate with minimal impact on losses.

Cavity flameholding in an optical axisymmetric scramjet in Mach 4.5 flows

The dynamics of an optical axisymmetric scramjet, preferred for generic combustion and flameholding studies without the corner boundary-layer effects that distort the freestream and flame propagation in rectangular scramjets, is investigated at Mach 4.5. Ethylene fuel is injected into supersonic flow via sixteen sonic fuel nozzles equally spaced in a circumferential direction and inclined at 45° to the freestream. The gaseous fuel is auto-ignited by high-enthalpy flows that are compressed and decelerated by shockwaves and boundary layers while passing through the scramjet inlet and isolator. An axisymmetric cavity located downstream of fuel injectors achieves flame holding by providing slow flow recirculation regions and inducing shockwaves at cavity leading edge and ramp to redirect and recompress flows. Flow and flame behaviors are characterized by high speed flame chemiluminescence imaging and static pressure measurement, while the detailed flame structures are resolved by instantaneous ground-state hydroxyl radical (OH) distributions using planar laser-induced fluorescence (PLIF). Numerical simulation is employed to aid the inlet/isolator design to avoid unexpected unstart introduced by shock-boundary layer interactions in the shock train region. The Mach number profiles in the radial direction are measured using a Pitot probe at the combustor exit. We demonstrate supersonic flameholding with the presence of a cavity at ethylene fueled conditions where stable combustion is achieved without scramjet unstart under both mass and heat loading. The heat addition from the cavity-stabilized flame mainly reduces the flow Mach number in the near-wall flow region. In the core flow region, the flow Mach number is decreased by the jet- and cavity-induced shockwaves and minimally affected by heat addition.

Goal-oriented error control of stochastic system approximations using metric-based anisotropic adaptations

The simulation of complex nonlinear engineering systems such as compressible fluid flows may be targeted to make more efficient and accurate the approximation of a specific (scalar) quantity of interest of the system. Putting aside modeling error and parametric uncertainty, this may be achieved by combining goal-oriented error estimates and adaptive anisotropic spatial mesh refinements. To this end, an elegant and efficient framework is the one of (Riemannian) metric-based adaptation where a goal-based a priori error estimation is used as indicator for adaptivity.This work proposes a novel extension of this approach to the case of aforementioned system approximations bearing a stochastic component. In this case, an optimization problem leading to the best control of the distinct sources of errors is formulated in the continuous framework of the Riemannian metric space. Algorithmic developments are also presented in order to quantify and adaptively adjust the error components in the deterministic and stochastic approximation spaces.The capability of the proposed method is tested on various problems including a supersonic scramjet inlet subject to geometrical and operational parametric uncertainties. It is demonstrated to accurately capture discontinuous features of stochastic compressible flows impacting pressure-related quantities of interest, while balancing computational budget and refinements in both spaces.