Inflow Mach Number Research Articles

Effects of air properties on flow filed characteristics of inflatable reentry decelerator

Air properties are very important for predicting the flow filed characteristics of aircraft. At present, Sutherland formula is always used to predict the flow filed characteristics of inflatable reentry decelerator. However, high temperature can be observed in high speed reentry flow around inflatable reentry decelerator, thus the Sutherland formula is not suitable for reentry flow simulation of inflatable reentry decelerator. In order to investigate the effect of the air properties on the flow filed characteristics of inflatable reentry decelerator, and provides the theoretical references for the inflatable reentry decelerator design, the Gupta fitting formula and Sutherland formula are used to describe the air properties, respectively, for studying the flow filed characteristics of inflatable reentry decelerator. The results show that the reentry flow temperature with Gupta fitting formula used is lower than that with Sutherland formula used, and the temperature difference is more obvious in the high inflow Mach number and temperature cases. The results in the windward of inflatable reentry decelerator show that the Sutherland formula underestimates the heat flux and pressure, and the heat flux difference is much more obvious. Such as case 5, the peak heat flux obtained by Gupta fitting formula is by 29% higher than that obtained by Sutherland formula.

Suitability of helium gas as surrogate fuel for hydrogen in H2-Air non-reactive supersonic mixing studies

Owing to the safety concerns associated with the usage of hydrogen (H 2 ) fuel in conducting experiments, the current study investigates the suitability of using helium (He) gas as a surrogate fuel for H 2 for supersonic mixing studies. The study adopts a steady state computational approach to compare the mixing performance between the two injection cases, where a wall-based pylon-cavity aided flameholder configuration is used for the investigation. A common fuel injection location at the cavity floor is used for He and H 2 injection cases with an injector hole diameter of 1 mm. An inflow Mach number of 2.2 with sonic transverse fuel injection is kept the same for all the cases. A steady Reynolds-averaged Navier–Stokes (RANS) based computational model for compressed real gas is used for the investigation. The governing equations are solved by an implicit, second-order upwind solver with a two-equation Menter's Shear Stress Transport (SST) turbulence model. The study shows a variation in near field mixing performance between He and H 2 injection cases due to the difference in the molecular mass. As a result, a 10%–11% variation in the mixing efficiency is observed between the two cases. This makes He not suitable for micro-level mixing studies as a surrogate to H 2 . However, it is possible to closely predict the trend in global mixing performance parameters as in the H 2 injection case. • H2 injection shows an enhanced near field mixing than He injection. • Trend in global mixing performance between He and H2 injection are similar. • He cannot be used as a surrogate fuel for H2 in micro-level mixing studies.

An Ultra-Fast TSP on a CNT Heating Layer for Unsteady Temperature and Heat Flux Measurements in Subsonic Flows.

In this paper, the authors demonstrate the application of a modified Ru(phen)-based temperature-sensitive paint which was originally developed for the evaluation of unsteady aero-thermodynamic phenomena in high Mach number but short duration experiments. In the present work, the modified TSP with a temperature sensitivity of up to −5.6%/K was applied in a low Mach number long-duration test case in a low-pressure environment. For the demonstration of the paint’s performance, a flat plate with a mounted cylinder was set up in the High-Speed Cascade Wind Tunnel (HGK). The test case was designed to generate vortex shedding frequencies up to 4300 Hz which were sampled using a high-speed camera at 40 kHz frame rate to resolve unsteady surface temperature fields for potential heat-transfer estimations. The experiments were carried out at reduced ambient pressure of = 13.8 kPa for three inflow Mach numbers being . In order to enable the resolution of very low temperature fluctuations down to the noise floor of K with high spatial and temporal resolution, the flat plate model was equipped with a sprayable carbon nanotube (CNT) heating layer. This constellation, together with the thermal sensors incorporated in the model, allowed for the calculation of a quasi-heat-transfer coefficient from the surface temperature fields. Besides the results of the experiments, the paper highlights the properties of the modified TSP as well as the methodology.

Onset of Oblique Detonation Waves for a Cavity-Based Wedge

To shorten the ignition length of the wedge-induced oblique detonation wave, a geometric setup in which the leading and aft wedges are connected by a cavity is proposed. High-resolution viscous computations to solve the reactive Reynolds-averaged Navier–Stokes equations have been performed, which examine the onset and evolution of combustion and flow configuration with premixed mixtures of inflow Mach numbers ranging from 4.0 to 6.0. The unreactive flowfield of the cavity-based wedge is analogous to that of a sole cavity with the exclusion of shock/shock interaction downstream. For and 5.0, no sustainable oblique detonation wave is achieved, and the flame holds in the cavity and near-wall area of the aft wedge; however, a short-lived detonation wave triggered by the shock focuses is observed during the combustion procedure with . For , nearly direct detonation initiation of postshock occurs, and an oblique detonation wave is established above the fore edge of the cavity. Unexpectedly, a delayed ignition downstream contributes to the development of a recirculation zone, which results in a detached oblique detonation wave for .

Numerical investigation of free oblique detonation wave induced by non-intrusive energy deposition

To validate the concept of free oblique detonation waves (ODWs) induced by non-intrusive energy deposition applied in ODW engines (ODWEs), numerical simulations are performed by modeling the laser spark as a high-temperature and high-pressure zone embedding in the high-speed incoming flow and by solving the non-dimensional reactive Euler equations with one-step chemistry. The initiation process, flow structures, and self-adapted characteristics of such kinds of free ODWs in an open space are discussed with varying dimensionless source energy and inflow Mach number. The results show that free ODWs can be initiated successfully by local non-intrusive energy deposition. In particular, four ignition regimes, namely, the subcritical regime, the critical I regime, the critical II regime, and the supercritical regime, which are analogous to the one-dimensional direct initiation of detonation, are distinguished by varying the dimensionless source energy. However, the typical energy required by each ignition regime is higher than that in the one-dimensional case due to multi-dimensional energy transmission and instabilities. Moreover, the dimensionless source energy should be moderate to obtain a stable free ODW flow field because extremely low source energy leads to ignition failure, while extremely high source energy causes premature ignition. When the inflow Mach number varies, it is found that free ODWs can adjust automatically to a nearly Chapman–Jouguet state, indicating the self-sustaining nature of free detonation waves. Due to this feature, ODWE performance will be relatively stable despite the change in inflow Mach numbers if free ODWs are utilized in the combustor.

Aerothermodynamic characteristics of hypersonic curved compression ramp flows with bistable states

The aerothermodynamic characteristics (including wall friction Cf, pressure pw, and heat flux St) in curved compression ramp (CCR) flows with bistable states are studied numerically and theoretically. Direct numerical simulations of bistable states of separation/attachment induced by variation of the turning angle ϕ, as well as the influence of inflow Mach number and wall temperature on dual-solution interval, are carried out. Distributions of wall friction, pressure and heat flux are analyzed. Further, the emergence of the first and second minima of Cf in the separation bubble is interpreted, which are dominated by the adverse pressure gradient induced by separation and reattachment shock, respectively. The present results and analysis indicate that the reversed-flow singularity of F. T. Smith (Proc. R. Soc. London. A: Math. Phys. Sci. 420, 21–52 (1988)] is less likely to occur in CCR flows. The prediction of pressure peak ppk of separation states confirms the model based on the minimum viscous dissipation theorem [Hu et al., Phys. Fluids 32(10), 101702 (2020)]. While the pressure overshoot pos can be analyzed by shock-polars with pressure match of compression and expansion process. The correlation between heat flux peak and pressure peak of both separation and attachment states is also discussed in terms of the classical power relation.

Efficient diffuser design for plasma wind tunnels with a large blockage model

The operation of plasma wind tunnels requires an excessive pressure ratio owing to the excess total pressure loss in the hypersonic environment. This loss increases when a large blockage model is mounted on the test section. Therefore, designing a diffuser that can alleviate this total pressure loss by compressing the hypersonic flow efficiently is essential. However, the flow inside a plasma wind tunnel diffuser exhibits a low Reynolds number and a high Mach number, rendering the prediction of the flow and performance of a diffuser difficult. Furthermore, limited literature exists on the design of plasma wind tunnel diffusers. In the study, we propose an efficient plasma wind tunnel diffuser capable of testing despite a large blockage model. We carried out flow analysis and a parametric study on the shape of the diffuser and numerically investigated the influence of a large blockage model. Additionally, a performance analysis was conducted to estimate the pressure ratio required for the operation of the wind tunnel. Throughout the study, the newly proposed parameter, mass flow rate weighted average of the total pressure, was identified as the primary observation value because it was found helpful for improving diffuser efficiency in the design stage. Furthermore, several design considerations are recommended for the enhancement of plasma wind tunnel diffusers. The inflow Mach number and total pressure at the nozzle exit were 7 and 21.9 bar, respectively, corresponding to conditions of MW class large plasma wind tunnels. With limited related studies in current literature, we expect this study to serve as a reference when designing a plasma wind tunnel diffuser.

THEORICAL STUDY ON PROPULSIVE PERFORMANCE OF OBLIQUE DETONATION ENGINE

Detonation combustion is characterized by the high thermodynamic efficiency and fast heat release. Benefitting from these potential advantages, an oblique detonation wave (ODW) is introduced into the combustion chamber and oblique detonation engine (ODE) plays an important role in hypersonic air-breathing propulsion systems. Previous studies mainly focused on the initiation structures, standing features and wave systems of oblique detonation, but the global analysis of ODE propulsive performance is still absent at the macro-level. In this paper, the flow and combustion processes of an ODE are decomposed into four basic modules, named as inlet model, mixing model, combustion mode and nozzle model, respectively. We solve these four basic flow processes using theoretical methods and propose a systematically theoretical approach that can be used to predict the ODE propulsion performance. On the basis of previous ODW initiation structures and waves systems, four different combustion modes, i.e., over-driven ODW, Chapman-Jouguet ODW, over-driven normal detonation wave and oblique shock-induced constant-volume combustion, are chosen to describe the heat release processes of combustible mixture in the ODE combustor. The effects of different combustion modes on fuel specific impulse of the ODE are also analyzed. In addition, the influence mechanisms of inflow parameters, combustor parameters and intake-exhaust parameters on the thrust performance of ODE are also obtained, and the results show that the major factor of fuel specific impulse of an ODE consists mainly of the inflow Mach number and the expansion ratio of engine nozzle. Finally, combined with precious detonation research results, such as the standing features and initiation structures of oblique detonation in a confined space, the preliminary design direction of oblique detonation engine are proposed, which mainly involve some constrained conditions, such as geometrical constraints, inflow velocity limitations and stability ranges of a detonation wave in ODE combustor.

The origin of a universal filament width in molecular clouds

ABSTRACT Filamentary structures identified in far-infrared observations of molecular clouds are typically found to have full widths at half-maximum $\sim 0.1\, {\rm pc}$. However, the physical explanation for this phenomenon is currently uncertain. We use hydrodynamic simulations of cylindrically symmetric converging flows to show that the full width at half-maximum of the resulting filament’s surface density profile, $\rm{\small FWHM}{_\Sigma }$, is closely related to the location of the accretion shock, where the inflow meets the boundary of the filament. For inflow Mach number, ${\cal M}$, between 1 and 5, filament $\rm{\small FWHM}{_\Sigma }$s fall in the range $0.03\, {\rm pc}\lesssim \rm{\small FWHM}{_\Sigma }\lesssim 0.3\, {\rm pc}$, with higher ${\cal M}$ resulting in narrower filaments. A large sample of filaments, seen at different evolutionary stages and with different values of ${\cal M}$, naturally results in a peaked distribution of $\rm{\small FWHM}{_\Sigma }$s similar in shape to that obtained from far-infrared observations of molecular clouds. However, unless the converging flows are limited to ${\cal M} \lesssim 3$, the peak of the distribution of $\rm{\small FWHM}{_\Sigma }$s is below the observed $\sim 0.1 \, {\rm pc}$.

A Parametric Blade Design Method for High-Speed Axial Compressor

The blade geometry design method is an important tool to design high performance axial compressors, expected to have large design space while limiting the quantity of design variables to a suitable level for usability. However, the large design space tends to increase the quantity of the design variables. To solve this problem, this paper utilizes the normalization and subsection techniques to develop a geometry design method featuring flexibility and local adjustability with limited design variables for usability. Firstly, the blade geometry parameters are defined by using the normalization technique. Then, the normalized camber angle f1(x) and thickness f2(x) functions are proposed with subsection techniques used to improve the design flexibility. The setting of adjustable coefficients acquires the local adjustability of blade geometry. Considering the usability, most of the design parameters have clear, intuitive meanings to make the method easy to use. To test this developed geometry design method, it is applied in the design of a transonic, two flow-path axial fan component for an aero engine. Numerical simulations indicate that the designed transonic axial fan system achieves good efficiency above 0.90 for the entire main-flow characteristic and above 0.865 for the bypass flow characteristic, while possessing a sufficiently stable operation range. This indicates that the developed design method has a large design space for containing the good performance compressor blade of different inflow Mach numbers, which is a useful platform for axial-flow compressor blade design.

Experimental study of hysteresis and catastrophe in a cavity-based scramjet combustor

Hystereses and catastrophes were experimentally investigated in a cavity-based scramjet combustor. The inflow Mach number was 3.0. Fuel Equivalence Ratio (ER) was continuously regulated with multi-steps to explore influences of historical regulation directions on combustion states. Two divided hysteresis loops with catastrophes were observed. By 1-D flow estimations, the first loop occurred with shock-free/separated scramjet mode transitions, while the second kept in the separated scramjet mode. This breaks through the traditional knowledge that hysteresis and catastrophe were certainly related to ramjet/scramjet mode transitions. The first hysteresis and catastrophes were attributed to flame stabilization mode transitions between the cavity shear-layer stabilized and the jet-wake stabilized, with flow separation establishment/vanishment upstream the cavities. The obvious variations of flame and shock/separation structures meant large wall-pressure changes in the expansive duct, and generated obvious thrust catastrophes. Besides, transition ER and catastrophe were larger in historical ER-increasing path because combustion efficiency became obviously larger as flow separation established. Difference of critical transition ERs meant the first hysteresis. The second hysteresis and catastrophes in the jet-wake stabilized mode were attributed to flame/shock interaction mode transitions between the flame/shock weak interaction mode and intensive interaction mode. Each transition caused slightly stronger/weaker flame interacting with slightly larger/smaller flow separation, which meant small wall-pressure changes in the expansive duct, and thus thrust catastrophe was unobvious. Hysteresis occurred as the critical transition ER was slightly higher in historical ER-increasing path because of slightly lower combustion efficiency under slightly smaller separation.

Numerical Investigation on the Effect of Inflow Mach Numbers on the Combustion Characteristics of a Typical Cavity-Based Supersonic Combustor

The air-breathing engines, commonly known as Supersonic Combustor Ramjet (SCRAMJET) engines, are one of the most prominent technologies among researchers due to their high thrust-to-weight ratio. The researchers are constantly making efforts for improved performance of the combustor under the required boundary conditions. The present working computational model studies a hydrogen-fueled parallel cavity scramjet combustor to recognize the complex flow field characteristics and performance of the combustor in Ansys 15.0. The computational model developed is a replica of an experiment conducted in China which slightly modified the boundary conditions. The standard two-equation K- ε turbulence model and Reynolds averaged Navier Stokes (RANS) equation with finite-rate/eddy dissipation species reaction model are used to simulate the problem. The validation of the present model is achieved by comparing the results with already available experimental data in conformity with the literature. The results of the simulations are in satisfactory accord with the experimental data and images. Furthermore, to achieve the stated objective, different incoming Mach numbers, namely, 2.25, 2.52, and 2.75, are considered for a more clear understanding of variables that affects the characteristics of the flow field. The temperature, Mach number, density pressure, and H2O mass fraction contours were studied to facilitate proper understanding. The maximum temperature rise observed is 2711.467 K for M = 2.25. Additionally, the performance parameters, namely, combustion and mixing efficiencies, are also studied. The maximum combustion and mixing efficiencies are 87.47% and 98.15% for M = 2.25 and 2.75, respectively.

A novel experimental method to the internal thrust of rocket-based combined-cycle engine

• The method can be used to study high altitude performance of the RBCC engine. • There is a great moment to the engine when the rocket turns on. • The air mass flow rate is the most important factor to the RBCC engine. The rocket-based combined-cycle engine is considered as a promising scheme to achieve affordable and reusable space transportation. Under the ejector mode, the air mass flow rate is limited by aerodynamic chocking in the mixer of the engine. In other words, the air mass flow rate is determined by the downstream mixing. However for the requirement of critical air mass flow rate, the air stream supplying and the internal thrust measurement are difficult to conduct in the conventional ground experimental facilities. And more importantly, the critical thrust performance is also hard to obtain. A novel experimental facility is introduced in which the air stream is entrained from the ambient atmosphere. The inlet throat is large enough to make sure the inflow air is throttled by aerodynamic chocking in the mixer. In addition, a supersonic ejector is used to simulate the pressure environment at a high altitude. The internal thrust under different flight conditions can be calculated through the measured parameters in the experiments. The results show that the novel experimental facility can be used to study the thrust performance of rocket-based combined-cycle engine in ejector mode. The test precision is acceptable and the thrust difference between the test and the simulation is less than 8%. With the increase of the inflow Mach number, the engine internal drag under the cold flow condition rises as well as the specific impulse augmentation. The specific impulse augmentation could be 16.5% when the entrainment ratio is 1.88 at a simulated condition of H6.8 km / Mach 1.2.

Design and validation of a two-dimensional variable geometry inlet

Abstract The design of a two-dimensional variable geometry inlet which applied to a tandem type turbine-based combined cycle (TBCC) propulsion system was investigated in the present paper through three-dimensional simulations and wind tunnel tests. The operation Mach number range was between 0 and 3. A multi-ramp geometry scheme was adopted to achieve acceptable performance at different inflow Mach number. The first ramp angle was fixed whilst the angles of the second and the third ramps were variable at different inflow Mach numbers. The Mach numbers at throat region were maintained between 1.3 and 1.5 at different inflow Mach numbers according to this variable geometry scheme. A fixed geometry rectangular-to-circular shape diffuser was adopted to improve aerodynamic performance of the inlet. Three-dimensional numerical simulations were carried out between Ma1.5 and Ma3.0. The results indicated that good aerodynamic performance can be achieved at different inflow speed. At the design point, total pressure recovery of the inlet was 0.66 at critical condition. Wind tunnel validation experiment tests were conducted at Ma2.0, showing the movement of terminal shock wave from downstream to upstream as the back pressure increased. The inlet operated at supercritical, critical and subsonic conditions at different back pressure.

Design and evaluation of a conical hypersonic vehicle with an overturned aerodynamic layout

A conical hypersonic vehicle with an overturned aerodynamic layout was accomplished and its aerodynamic performances were evaluated. In this work, the aerodynamic design methodology was addressed firstly, and then the conical vehicle integrated with an overturned inlet was obtained. The mass flow capture performance and the total pressure recovery performance were evaluated under design and off-design conditions. Results show that the mass flow capture increases and the total pressure recovery decreases as the inflow Mach number increases. The total pressure recovery decreases when the inflow Mach number increases. After that, the self-starting performance of the inlet was obtained by simulating an accelerating process. The initial inlet self-starts at Mach 5.2 on attack angle 6∘ situation. Furthermore, an improvement of the inlet self-starting performance was achieved with keeping the designed mass flow capture almost constant. By altering the 3-D shape of the inlet leading edge, the inlet self-starting Mach number decreases from Mach 5.2 to Mach 4.6 on attack angle 6∘ situation.

Numerical Analysis and Wind Tunnel Validation of Low-Temperature Ablators undergoing Shape Change

Ground testing of ablative materials aims at providing critical data on the material behavior under hypersonic reentry conditions. This is normally done in plasma wind tunnel facilities. However, non-negligible technical challenges are faced in order to duplicate the real flight conditions, such as inducing the recession of space-relevant ablative materials, which requires sufficiently high inflow total enthalpies, and/or reproducing the actual hypersonic flow velocity, which requires sufficiently high inflow Mach numbers. Often, ground facilities which are providing one requirement are lacking the other one and vice-versa. A possible solution is to use low-temperature ablators in continuous hypersonic blow-down tunnels, where aerodynamic and ablative tests with considerable shape change effects may be performed under reasonably low total temperature conditions and with affordable test durations. These substances are readily available, and they sublimate or ablate in a fashion that can be described fairly accurately by theory. This work has the objective to numerically characterize the shape change of such materials in hypersonic conditions, concurrently providing a validation against literature data and from a dedicated experimental ground test campaign. The numerical procedure relies on ad-hoc mesh generation/evolution strategies taking into account the material shape change, and is based on subsequent steady-state Computational Fluid Dynamics (CFD) computations coupled with a customizable gas-surface interaction wall boundary condition. Preliminary numerical simulations helped the design of the experiments to be carried out in the von Karman Institute (VKI) H-3 hypersonic wind tunnel, in particular for the identification of capsule geometry and size in order to maximize the shape change caused by ablation. Subsequently, camphor is identified as the most suitable low-temperature ablator to be used in the experimental campaign after a thorough analysis of its surface reaction thermodynamics and kinetics. Results from the CFD approach are first compared with a literature experimental test case and then with those of the previously designed experiments, featuring a camphor sub-scale capsule, underlying advantages and limits of the numerical procedure adopted. The obtained numerical and experimental results underline how it is possible to obtain a relevant shape change for relatively small exposure times by using low-temperature ablators in continuous hypersonic blow-down wind tunnels. Hence, results from this work can be used to support the design and sizing of the actual heat shield and the analysis of the capsule’s aerodynamics and stability, accounting for shape change effects, by establishing an appropriate similitude between in-flight and on-ground conditions.

Numerical investigation on the extrusive and intrusive subcavity types and their location on the primary recirculation zone for the supersonic turbulent flow through cavity type flameholders

Supersonic turbulent non-reacting flow through a cavity-type flameholder employed in scramjet engines with and without subcavities is numerically investigated. The governing equations describing the flow are solved by employing a density-based solver i.e., rhoCentralFoam which is an open-source computational fluid dynamics (CFD) code in OpenFOAM. The turbulence is modeled using the two-equation k-ω SST (shear stress transport) turbulence model. In the present study implication of subcavity types (i.e., extrusive and intrusive) on recirculation patterns and their strength at different subcavity aspect ratios (l/d) and subcavity locations at different inflow Mach numbers is investigated. A significant increase in the strength of the primary recirculation zone is observed for modified rectangle cavity 2 (MRC2) compared to the base rectangle cavity (BRC) and modified rectangle cavity 1 (MRC1). Results show a significant decrease in the size of the secondary recirculation zone for modified angle cavity 2 (MAC2) for all the subcavity aspect ratios and subcavity leading-edge distances (SLDs) at Mach 2. Results also indicate that at a fixed Mach number, the strength of the primary recirculation decreases with the increase in base cavity aspect ratio (L/D) for all the cavity types. Results show a substantial decrease in the velocity magnitude of cavity fluid with the decrease in cavity aft wall angle from 90° to 30°, which shows that the strength of the primary recirculation zone for BRC and MRC2 is higher compared to base angle cavity (BAC) and MAC2. Results show a 20.4776 % and 7.3822 % increase in the peak values of streamwise and normal velocities for MRC2 compared to BRC and MRC1 at Mach 2, and the corresponding increase in velocities for Mach 3 are 14.7979% and 5.6693% respectively for all the aspect rations and SLDs. Present results are validated with the experimental results available in the literature.

Effects of fuel injection speed on supersonic combustion using separation-resistant struts

This paper describes the improvement in combustion efficiency achieved using streamwise vorticity and fuel injections when using hypermixer struts at high Mach numbers. The effects of fuel injection conditions on combustion were investigated using numerical simulations of chemical reactive flows with a detailed reaction mechanism. Hydrogen fuel was injected from the trailing edges of a separation-resistant strut. Fuel injection at sonic speeds led to an enhancement in efficiency with an increased injection angle at Mach number 2.5, but this improvement was not observed at Mach number 3.5. It was found that hydrogen accumulated in the vortex core because of the strong circulation generated from the strut at high Mach numbers. With fuel injection at speeds exceeding the sonic speed, the efficiency increased, and it improved further with an increasing injection angle. From the results, an important indicator for combustion enhancement was derived from the streamwise vortex circulation, fuel injection velocity, and inflow Mach number. To increase the combustion efficiency using the strut, the indicator value should exceed a threshold. This is crucial for ensuring effective fuel consumption during combustion. These findings provide useful insights into a good design of hypermixer struts in supersonic combustion.

Role of corner flow separation in unsteady dynamics of hypersonic flow over a double wedge geometry

This study investigates the origin and sustenance of self-induced oscillations of shock structures in a hypersonic flow over a double wedge configuration. Previously, various researchers have considered the double wedge flow configuration for inviscid flow with variations of different inflows as well as geometric parameters such as inflow Mach number (M∞), wedge angles, and wedge lengths. Few recently published articles reveal an unsteady flow physics involved with the hypersonic viscous flow for double wedge configuration with large second wedge angles. However, the reason for such self-sustained flow oscillations is not completely clear. The present work seeks out to investigate the origin of such oscillations in a low enthalpy hypersonic flow with different aft-wedge angles and wedge length ratios. In the current study, viscous flow over a double wedge at M∞ = 7 and fore-wedge angle of 30° is considered. An improved version of rhoCentralFoam solver in OpenFOAM is used to investigate the double wedge flow over different grid resolutions in the separation region and shock–shock interaction region. This study corroborates the observation from the previous literature with an improvement in the range of parameters which results in a self-sustained periodic oscillation. The present study also suggests that the unsteadiness becomes possible when the incidence shock is in the proximity of the aft-wedge expansion corner as a consequence of different wedge length ratios (L1/L2) or aft-wedge angles (θ2). Flow can still be steady at a large aft-wedge angle if the incidence shock is far ahead of the aft-wedge expansion corner.

On the Importance of Transition Control at Transonic Compressor Blades

AbstractThe flow through a transonic compressor cascade is characterized by high unsteadiness and a high loss level. This results from the shock waves in the blade cascade and their interaction with the blade suction side boundary layer. In the case of a laminar shock wave boundary layer interaction, the loss level is higher due to the occurrence of a laminar separation bubble below the shock wave compared to the shock wave interaction with a turbulent boundary layer. In addition, the oscillation of the shock position in both cases influences the working range concerning the point of stall onset as well as leading to an unsteady interaction with the blade called buffeting. The reduction of losses and of unsteadiness in the shock wave oscillation, connected to a decrease of the blade buffeting effect, is the aim of the current investigation. Therefore, experimental investigations using a roughness patch as well as air jet vortex generators in order to control the transition in a transonic compressor cascade have been conducted at the transonic cascade wind tunnel of the German Aerospace Center (DLR) at Cologne. At an inflow Mach number of 1.21, a loss reduction for both transition control cases is achieved. In spite of a nearly uninfluenced fluctuation range of the passage shock wave compared to the reference cascade, the oscillation spectra of the transition control cases show a reduction of the shock movement amplitude at a frequency below 500 Hz and above 1 kHz. In the closing section of the paper, a detailed discussion on the reasons for the resulting flow behavior based on particle image velocimetry and high-speed shadowgraphy data is given. The resulting conclusion of the study is that the consideration of transition control at transonic compressor blades is very important in order to reduce losses and flow unsteadiness that directly influences blade buffeting and the numerical prediction quality of the stall onset.