Cowl Shock Research Articles

Design and aerodynamic characteristics of variable-geometry hypersonic inlet based on shock wave elimination

To solve the problem that the control of shock wave elimination is weakened under off-design conditions, the design concept of a variable-geometry inlet scheme that combines the variable-geometry cowl (translating and diagonalizing) with regulating shock wave elimination is introduced in this paper. The variable-geometry inlet is designed by the theories of oblique shock wave and isentropic wave as well as the Oswatitsch theory. Regulatory law of the variable-geometry cowl based on shock wave elimination is obtained by the geometric relationships between cowl compression angle, cowl shock wave angle, and optimal control point or range. Numerical simulations are conducted to investigate flow field characteristics, control mechanism, and working performance of the inlet. Results reveal that expansion waves have a significant impact on the cowl shock wave and boundary layer interaction, and flow separation. Furthermore, variable-geometry inlet with translating and diagonalizing cowl based on the regulation of shock wave elimination effectively controls and even completely eliminates the flow separation. In terms of inlet performance, the total pressure loss of the variable-geometry inlet decreases such that the total pressure recovery coefficients of the translating cowl and diagonalizing cowl inlets are increased by maximum values of 3.39 % and 9.97 %, respectively. However, the mass flow coefficient of translating cowl inlet decreases, whereas that of the diagonalizing cowl inlet is equivalent to that of the fixed-geometry inlet. The working range can be widened by changing the internal contract ratio of the inlet through translating or diagonalizing the cowl. The results confirm that the scheme of variable geometry inlet with diagonalizing cowl is practicable and reliable and has important guiding significance and value for inlet design.

Transient Flow Evolution of a Hypersonic Inlet/Isolator with Incoming Windshear

In this paper, a novel flow perturbation model meant to investigate the effects of incoming wind shear on a hypersonic inlet/isolator is presented. This research focuses on the transient shock/boundary layer interaction and shock train flow evolution in a hypersonic inlet/isolator with an on-design Mach number of 6.0 under incoming wind shear at high altitudes, precisely at an altitude of 30 km with a magnitude speed of 80 m/s. Despite the low intensity of wind shear at high altitudes, the results reveal that wind shear significantly disrupts the inlet/isolator flowfield, affecting the shock wave/boundary layer interaction in the unthrottled state, which drives the separation bubble at the throat to move downstream and then upstream. Moreover, the flowfield behaves as a hysteresis phenomenon under the effect of wind shear, and the total pressure recovery coefficients at the throat and exit of the inlet/isolator increase by approximately 10% to 12%. Furthermore, this research focuses on investigating the impact of wind shear on the behavior of the shock train. Once the inlet/isolator is in a throttled state, wind shear severely impacts the motion of the shock train. When the downstream backpressure is 135 times the incoming pressure (p0), the shock train first moves upstream and gradually couples with a cowl shock wave/boundary layer interaction, resulting in a more significant separation at the throat, and then moves downstream and decouples from the separation bubble at the throat. However, if the downstream backpressure increases to 140 p0, the shock train enlarges the separation bubble, forcing the inlet/isolator to fall into the unstart state, and it cannot be restarted. These findings emphasize the need to consider wind shear effects in the design and operation of hypersonic inlet/isolator.

Similarity Analysis for Two-Dimensional Hypersonic Inlets Restart

The restart ability of inlets is significant to the robustness of hypersonic airbreathing vehicles. Classic starting theories based on the one-dimensional and inviscid flow assumption cannot predict the restart boundary accurately. To overcome the classic theories, the restart characteristics of two-dimensional contraction ducts in viscous flow are numerically studied. The restart boundaries for different configurations and flow conditions are determined, and the restart similarity laws are discovered. With the proposed dimensionless quantities (the dimensionless cowl length and the effective throat-entrance area ratio), the influences of multiple factors on restart are described simultaneously in a concise self-similar formulation. The restart characteristics can be described with any two variables in cowl angle, dimensionless cowl length, and effective area ratio. The restarts of two-dimensional inlets are classified into three types of the short-cowl, transitional, and long-cowl restarts according to the features of restart boundary lines and the flow structure. The three restart types are dominated by the competition between the cowl shock system and cowl tail expansion waves, the cowl shock system alone, and the complementation between the cowl shock system and throat choke, respectively. The wall pressure criteria for restart are correlated with the dimensionless cowl length and the characteristic separation scale at the restart critical state. It is observed that Kantrowitz limit is not generalizable for restart boundaries.

Grid Adaptive Technique for Simulation of Scramjet Intake-Isolator at Hypersonic Speeds

Hypersonic intake is one of the major components of Scramjet engine. It compresses the incoming hypersonic flow through a series of oblique shocks as the flow passes through intake-isolator section before entering the combustion chamber, which is essential for efficient combustion. The shocks generated inside in the intake interacts with boundary layer following shock boundary layer interaction and flow separation. The separated flow blocks the flow capture area such that engine expresses unstarting phenomenon. Understanding and mitigating such flow phenomenon is a challenging task. With respect to hypersonic speeds the experimental facilities are very limited. The only alternative to solve this problem is Computational Fluid Dynamics because of its capabilities. But validation of CFD results with analytical or experimental is the foremost prerequisite to chase computational analysis. Mostly at high speeds the precision of CFD results rest on the type of grid, number of elements and turbulence model used. So, in this paper, computational analysis of hypersonic intake is carried out through designed conditions to ensure the correct CFD process is used by varying number of elements in fluid domain by grid adaptive technique using ANSYS Fluent and satisfying Y+ parameter. The domain is analysed with various turbulence models and among them SST has predicted all the flow characteristics of scramjet intake-isolator at hypersonic speeds like separation bubble, shock reattachment, cowl shock etc similar to experimental results with the help of grid adaptive technique. So, grid adaptive technique is also proposed for simulation of scramjet intake at off-design conditions.

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.

Computation of viscous flow characteristics of a supersonic intake using eddy-viscosity turbulence models

Results of numerical investigation of flow and performance characteristics of a rectangular supersonic intake experimentally tested at Mach number of 4.03 are presented. The viscous flowfield has been obtained by solving Favre averaged Navier–Stokes equations with two eddy viscosity turbulence models, namely, Mentor’s baseline (BSL) k-ω and shear stress transport (SST) k-ω models. In addition, the effect of wall temperature conditions as well as several configurations of cowl and isolator are simulated to validate the computational methodology. The analysis has been carried out at a freestream Mach number of 4.03 and 0° angle-of-attack, corresponding to a unit Reynolds number of about 6.9×107 m−1. The predicted boundary-layer and the wall static pressure distribution using SST k-ω model shows better agreement with the experimental data compared to the BSL model and hence indicates superior capability to capture the onset and the amount of flow separation due to SWBLIs. Moreover, the terminal shock structure at different operating modes as well as the “unstart because of SWBLIs” is predicted quite accurately using the model. The wall temperature condition is found to have a strong effect on the amount of flow separation at the cowl and the subsequent interactions. Further, the results show that the sidewalls have a marginal effect on the centerline wall static pressure, though a three-dimensional flowfield with two stream-wise vortices are observed close to it as a result of the swept SWBLIs of the cowl shock with the sidewall boundary-layer leading to a higher loss in performance.

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.

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.

Investigation of local unstart in a hypersonic scramjet intake at a Mach number of 6

Understanding the mechanism of local unstart in a mixed compression scramjet intake is vital for its stable off-design operation. In the present work, the local unstart behavior of a hypersonic scramjet intake is studied at a Mach number of 6, and a flight realistic Reynolds number of 8.4×106/m using a combination of experiments in a hypersonic wind tunnel and numerical computations. To study the effect of internal contraction ratio (ICR), experiments are performed on an intake model that is equipped with an all movable cowl in two modes: (a) cowl fixed at a particular ICR, and (b) cowl moved to a particular ICR during the steady test time of the wind tunnel. The intake exhibits started flow for ICR between 1.19 and 1.37, and locally unstarted flow for ICR ≥1.45. In the dynamic cowl experiments, the impingement of the forebody shock onto the inner cowl surface forms a leading-edge separation bubble, which moves upstream and detaches the cowl shock, leading to local unstart with subsonic spillage of mass near cowl lip. A shear layer emanating from the triple point of shock interaction divides this subsonic flow and the supersonic flow at the ramp side. The frequency content of the flow is different between the fixed and the dynamic cowl cases. Local unstart observed during fixed cowl experiments reveal low amplitude self-sustained oscillations around a frequency of 2.5 kHz. The dynamic mode decomposition (DMD) and proper orthogonal decomposition (POD) of numerical density contours reveal that the shear layer impingement on the cowl surface drives the self-sustained oscillations observed during local unstart.

Effects of Wall Ventilation on the Shock-wave/Viscous-Layer Interactions in a Mach 2.2 Intake

In order to achieve proficient combustion with the present technologies, the flow through an aircraft intake operating at supersonic and hypersonic Mach numbers must be decelerated to a low-subsonic level before entering the combustion chamber. High-speed intakes are generally designed to act as a flow compressor even in the absence of mechanical compressors. The reduction in flow velocity is essentially achieved by generating a series of oblique as well as normal shock waves in the external ramp region and also in the internal isolator region of the intake. Thus, these intakes are also referred to as mixed-compression intakes. Nevertheless, the benefits of shock-generated compression do not arise independently but with enormous losses because of the shockwave and boundary layer interactions (SBLIs). These interactions should be manipulated to minimize or alleviate the losses. In the present investigation a wall ventilation using a new cavity configuration (having a cross-section similar to a truncated rectangle with the top wall covered by a thin perforated surface is deployed underneath the cowl-shock impinging point of the Mach 2.2 mixed-compression intake. The intake is tested for four different contraction ratios of 1.16, 1.19, 1.22, and 1.25, with emphasis on the effect of porosity, which is varied at 10.6%, 15.7%, 18.8%, and 22.5%. The introduction of porosity on the surface covering the cavity has been proved to be beneficial in decreasing the wall static pressure substantially as compared to the plain intake. A maximum of approximately 24.2% in the reduction in pressure at the upstream proximal location of 0.48 L is achieved in the case of the wall-ventilated intake with 18.8% porosity, at the contraction ratio of 1.19. The Schlieren density field images confirm the efficacy of the 18.8% ventilation in stretching the shock trains and in decreasing the separation length. At the contraction ratios of 1.19, 1.22, and 1.25 (‘dual-mode’ contraction ratios), the controlled intakes with higher porosity reduce the pressure gradients across the shockwaves and thereby yields an ‘intake-start’ condition. However, for the uncontrolled intake, the ‘unstart’ condition emerges due to the formation of a normal shock at the cowl lip. Additionally, the cowl shock in the ‘unstart’ intake is shifted upstream because of higher downstream pressure.

Effect of cowl shock on restart characteristics of simple ramp type hypersonic inlets with thin boundary layers

The effect of cowl angle on the restart characteristics of simple ramp type hypersonic inlets was experimentally investigated in shock tunnel equipped with schlieren imagery and static pressure measurement. The cowl shock strength is found to be a key factor that determines the inlet restart and makes the restart contraction ratios significantly deviate from the Kantrowitz criterion. Stronger cowl shock tends to degrade the inlet restart capability by causing larger separation bubble and higher pressure loss during the restarting process. In particular, a sensitive range of the cowl angles, within which the restart contraction ratio decreases rapidly, was identified. A design concept of multiple noncoalesced cowl shocks was thus proposed and proven to significantly improve the inlet restart capability.

Control of Severe Shock-Wave/Boundary-Layer Interactions in Hypersonic Inlets

A theoretical analysis and numerical simulations are conducted to demonstrate the control capability of a splitter on the shock-wave/boundary-layer interactions in a generic hypersonic inlet with both a two-staged external compression and a strong cowl compression. The results show that, under the unstart condition, a well-designed splitter could suppress the massive cowl shock-induced separation, increase the fullness of the rampside boundary-layer profile at the isolator exit, and improve the inlet performance over the targeted operating Mach number range. With the splitter design, the severe shock-wave/boundary-layer interaction induced by the strong cowl shock is replaced by several weaker shock-wave/boundary-layer interactions with reduced shock intensity and boundary-layer thickness.

Investigation of hypersonic inlet pulse-starting characteristics at high Mach number

An unsteady viscous numerical simulation is performed to study the hypersonic inlet pulse-starting in a shock tunnel under the condition of inflow Mach number higher than the design value with a simplified nozzle startup wave system. A local unstart of hypersonic inlet is obtained with the initial backpressure and temperature of test section equal to the hypersonic freestream static pressure and static temperature, respectively. During the hypersonic inlet pulse-starting process, the nozzle startup wave system (including the primary shock, contact discontinuity and secondary shock) propagates downstream and interacts with the hypersonic inlet, the secondary shock induces boundary layer separation on the compression ramp. A low velocity flow induced by the separation bubble enters into the internal duct with the passage of time. It thickens the boundary layer and then the cowl shock induces a large separation bubble on the compression side. The shock induced by the large separation bubble impinges on the cowl and induces a small separation bubble, which moves upstream and induces the reflection of the external compression shock at the cowl transiting from regular reflection to Mach reflection. Mach reflection results in the shock propagating forward to cause a shock detachment at the cowl lip. The hypersonic inlet is not fully started and this phenomenon is referred to as “local unstart of inlet”. Moreover, it is found that the hypersonic inlet starts with lower initial backpressure or higher initial temperature of test section, and the hypersonic inlet unstarts with higher initial backpressure. The effect mechanism is analyzed with the help of numerical simulations and unsteady one-dimensional shock theory. As there are three operation modes of hypersonic inlet at high Mach number: start, local unstart and unstart. The results of the hypersonic inlet pulse-starting in a shock tunnel should be carefully treated.

Unstart Process of a Rectangular Hypersonic Inlet at Different Mach Numbers

A rectangular hypersonic inlet is tested at the low-speed mode ( 5), the design mode ( 6), and the overspeed mode ( 7) to enrich the understanding of hypersonic-inlet unstart. A flow plug is placed at the duct exit to simulate the combustion-induced high pressure and initiate the inlet unstart. High-speed schlieren imaging and time-resolved pressure measurements are used simultaneously to record the unsteady-flow structures and surface pressures of the unstart process. The results indicate that the unstart processes are basically the same for the low-speed mode and the design mode. An unstart phenomenon dominated by the ramp-side flow separation, which exhibits violent fluctuations along the entire duct with a low frequency, is observed at these two operating modes. While operating at the overspeed mode, the strong shear layer, induced by the Mach reflection of the coalesced external oblique shock, divides the internal flow into two parts. The cowl side is filled by fully subsonic flow after the compression of the detached normal shock, and the ramp side mainly contains supersonic flow downstream of the shear-layer-induced oblique shock. Therefore, a real-time adjustment of the detached cowl shock to the downstream high pressure is formed, keeping the entire flowfield of the inlet from oscillation in a much wider range of throttling ratio. With a further increase of the throttling ratio, an oscillation dominated by the cowl-side reversed flow occurs. The oscillation phenomenon exhibits relatively gentle fluctuations, and the characteristic frequencies are mainly determined by the shock train in front of the throat section.

Shock Impinging Effect on Ethylene Flameholding

The flameholding characteristics of ethylene in supersonic crossflows were investigated experimentally. Shock-tunnel experiments were conducted at a flow total enthalpy of , with a fuel–air equivalence ratio of 0.2. A contoured wedge model with a wall cavity and a cowl was used. To investigate the shock impinging effect on flameholding, different cowl locations were considered, varying from the upstream of the injection to the cavity ramp. Shadowgraph and CH* chemiluminescence optical images were obtained for each location. The experimental data showed that flameholding can be categorized as three distinct types of interaction. The first interaction, with the cowl shock impinging upstream of the fuel injection, showed a gradual quenching of CH* radical in the cavity during steady flow, and flameholding was not maintained. The second interaction, with the cowl shock impinging at the cavity ramp, showed flameholding in the shear layer above the cavity and downstream throughout. The third interaction, with the cowl shock impinging directly at the fuel injection, showed flameholding in the entire cavity and the downstream region. It was shown that, depending on the cowl location, ethylene flameholding could be achieved at a much lower enthalpy than a recent finding of flow total enthalpy of .

Validation of Three-Dimensional Simulation of Flow through Hypersonic Air-breathing Engine

<p>The flow path of a conceptual hypersonic air-breathing scramjet engine integrated with the vehicle (without combustion) has been simulated numerically using ANSYS CFX software with the SST turbulence model. The computations were performed for the free stream Mach number of 6 and angle-of-attack of 5°. A strong separation bubble was observed on the bodyside wall in the internal compression region where the reflected cowl shock impinges on body which in turn increases the static pressure substantially. The external-internal flow field of the hypersonic mixed compression intake, shock-boundary layer interactions, and the shock-shock interactions present in the internal compression region have qualitatively been obtained and analysed. The variation of centreline pressure along the bodyside wall close to the symmetry plane obtained from numerical simulation centreline has been compared with the experimentally measured data. It has been observed that the computed wall pressure matches fairly well with the measured values in the external ramp compression region, internal compression region and in the combustion chamber. The flow patterns and the pressure variations near the middle wall and the fuel injecting strut locations have also been analysed.</p><p><strong>Defence Science Journal, Vol. 65, No. 4, July 2015, pp. 272-278, DOI: http://dx.doi.org/10.14429/dsj.65.6979</strong></p>

Influence of Expansion Waves on Cowl Shock/Boundary Layer Interaction in Hypersonic Inlets

The cowl shock/boundary layer interaction is of great importance to the performance and operability of hypersonic inlets. However, it is usually affected by the expansion waves originated from the convex corner of the ramp surface (often called the shoulder), making the commonly used separation prediction relation unreliable. Therefore, the cowl shock/boundary layer interaction under the interference of the expansion waves is investigated by both theoretical analysis method and computational method in this paper. The results show that the expansion waves bring significant influence on the cowl shock/boundary layer interaction and four kinds of interaction processes exist when their relative position changes. When different interaction processes dominates the flowfield, the expansion waves may bring positive or negative effects on the shock/boundary layer interaction. In particular, while the cowl shock impinges near the shoulder, the interaction processes of shock–shock–expansion wave and shock–expansion wave–shock bring positive effects on the shock/boundary layer interaction and suppress the boundary layer separation.

A TIPSO algorithm assessment for aerothermodynamic optimization of hypersonic compression systems

A two-step improved particle swarm optimization (TIPSO) algorithm was recently proposed and used for the optimization of flexible satellite-control parameters. It was found to be much more stable and much less complex than other evolutionary algorithms. In this article the efficacy of the TIPSO algorithm is investigated for multidisciplinary optimization of aerothermodynamic parameters of performance, cowl deflection angle and shock–boundary layer separation on a cone-derived, wedge-integrated hypersonic (waverider) compression system. This algorithm uses an aggregate objective function. Optimization results from the TIPSO algorithm are compared with those obtained from a hybrid genetic algorithm, particle swarm optimization using an inertial weight approach and a multi-objective genetic algorithm. Since optimality of the forebody configuration is the basic requirement, the optimization variables selected are the isolator exit Mach number, static pressure ratio across the forebody–inlet configuration, cycle temperature ratio and cowl (inlet) deflection angle, which is the maximum deflection angle corresponding to the optimum oblique shock angle generated by the inlet cowl. Simulation results show that for the given case study only the TIPSO algorithm can locate the globally best aerothermodynamic–geometric parameters, compared with all the other optimization methods.

Experimental Investigation of a Hypersonic Inlet with and Without Sidewall Compression

The inlet of a Mach 8 cruise vehicle has been experimentally investigated in a hypersonic blow down wind tunnel with test times of up to 30s. Two test series have been carried out, the first one with a two-dimensional inlet configuration and the second one with the inlet featuring additional side-wall compression. The objective of the tests was to determine the performance of the inlet at on- and off-design conditions, i.e. at angle of attack and angle of yaw. Additionally, the starting behavior of the inlet has been determined, i.e. the contraction ratio at which the inlet starts. Results were interpreted from schlieren visualizations, IR thermography, static and Pitot pressure readings and mass flow measurements. The off-design investigation of both inlet configurations revealed the sensitivity of the inlet towards a cowl shock induced separation on the second compression ramp. If the cowl shock was not ensured to fall into the bleed duct behind the second ramp, no inlet flow could be established. The comparison of starting contraction ratios showed the advantage of the inlet with side-wall compression, i.e. the inlet with sidewall compression starts at higher contraction ratios.