Inflow Mach Number Research Articles

Shock-Induced Fan Cowl Separation During Aeroengine Windmilling at Diversion from Cruise

When a civil aircraft engine is operated at windmill during the cruise flight phase, there is supersonic flow acceleration around the leading edge of the fan cowl toward the external surface. The terminating normal shock wave can separate the turbulent boundary layer developing on this external surface. A series of experiments at a flight-relevant Reynolds number (1.2 million based on lip thickness) are performed in a quasi-two-dimensional wind tunnel rig to investigate the underlying flow physics. At a nominal inflow Mach number of 0.65 and a nacelle incidence angle of 4.5 deg, as the equivalent engine mass-flow rate is reduced, an increase in shock strength results in flow separation when the shock exceeds Mach 1.4. Over a 10% range in the notional engine mass-flow rate, the boundary layer developing on the external fan cowl thickens by a factor of three on the onset of separation. A reduction in the incoming Mach number from 0.65 to 0.60 weakens the shock wave and thus delays separation. An increase in surface roughness has no significant effect in situations where the boundary layer remains attached. However, for separated cases, an increased local roughness height causes a greater separation extent and a thicker boundary layer downstream of the shock wave.

Heat flux measurement at hypersonic turbulent separated flow field induced by the blunt fins

Abstract Hypersonic separated flow induced by the protrusions is one of the most concerning research studies in aerodynamics due to the complex flow structure and the significance of engineering. An experimental study of three-dimensional hyper-sonic separated flow induced by the blunt fins is investigated. The tests are performed with inflow Mach number 6. The main flow properties of the shock systems are described by analyzing Schlieren photographs. The heat flux data are obtained by using high precision platinum film sensors. The effects of the fin sweep angle are also presented. Results show that the maximum heat flux decreases, and the range of the separation zone quickly reduces as the swept angle increases. Low-frequency oscillations of separated flow fields are analyzed through instantaneous heat flux signals at typical measuring points. It is shown that characteristic frequencies of shock wave oscillation with no rear sweep angle is about 1 k Hz.

Transition of the flow type in the supersonic cavity controlled by the wall temperature

The wall temperature of the high-speed aircraft increases quickly under hypersonic/supersonic incoming flow, which will cause a significant change in the flow structure. To study the transition of the flow type in the supersonic cavity controlled by the wall temperature, numerical simulations are conducted. The cavity length-to-depth ratio (L/D) is varied from 10 to 15, and the wall temperature ranges from 300 K to 1300 K. The results indicate that the type of cavity flow with an L/D ratio of 13 transforms from a closed cavity flow to a transitional cavity flow, when the temperature reaches approximately 775 K. And the transitional temperature rises with the elevated total temperature of the incoming flow. Furthermore, the mechanism of the cavity flow change with wall temperature could be the competition between the recirculation zone and the shear layer in the cavity. The rising pressure with higher wall temperature in the recirculation zone weakens the downward development of the cavity shear layer, preventing it from hitting the cavity floor. As a result, the mass exchange of cavity lip surface, pressure distribution, total pressure recovery coefficient, and heat transfer distribution in the supersonic cavity change dramatically. The critical wall temperature also affected by the sidewall effects and the inflow Mach number.

Expansion wave-reinforced initiation of the oblique detonation wave

Efficient initiation of oblique detonation waves (ODWs) is crucial for optimizing the performance of oblique detonation wave engines. A novel approach is proposed for enhancing ODW initiation through expansion waves in this research. Validation of the expansion wave-reinforced initiation method is conducted via numerical simulations employing multi-species reactive Euler equations and a pressure-dependent reaction mechanism. Results demonstrate a significant reduction in the initiation length of ODWs with the addition of an expansion wave ahead of the wedge, contrasting with the absence of detonation wave initiation on a wedge lacking an expansion wave. A theoretical model, based on expansion wave and shock wave relations, along with constant volume combustion theory, elucidates the underlying mechanism of reinforcement. The model reveals that crossing the expansion wave elevates the fluid's Mach number and locally enlarges the flow deflection angle on the wedge surface, without altering the wedge's structure. Furthermore, post-shock temperature increases and pressure decreases compared to the wedge not encountering an expansion wave. The heightened temperature predominantly triggers ODW initiation, thus reinforcing the process. Theoretical analyses indicate the reinforcement's greater efficacy at lower inflow temperatures and lower inflow Mach numbers, suggesting the expansion wave's suitability for initiation in the early flight stages of an aircraft equipped with oblique detonation wave engines.

An experimental study on the combustion characteristics of the cavity-stabilized supersonic premixed flame

With the overwhelming trend of fast and economic aviation, it is becoming increasingly difficult for fuel to burn stably and efficiently in a constrained space and time period in the combustor of a hypersonic air-breathing vehicle. Conventional supersonic diffusion flame runs the risk of blowing out in the incoming flow at such high velocity. Upstream fuel injection of the flame holder is gaining popularity due to its prominent advantage in improving mixing efficiency, and this fuel injection method contributes to the organization of supersonic premixed combustion. In order to study the combustion characteristics of the supersonic premixed flame, corresponding experiments are carried out on a direct-connected scramjet platform with the inflow total temperature of 1700 K and the inflow Mach Number of 2.0. Gaseous ethylene with an equivalence ratio of 0.15 was injected transversely 670 mm upstream of a cavity. Optical measurements including laser-induced fluorescence on the hydroxyl radical (OH-LIF), CH*chemiluminescence photography, and high-speed photography were performed to investigate the steady-state and transient combustion characteristics of the cavity-stabilized supersonic premixed flame. From steady-state OH-LIF and CH* results, the flame could be divided into reactant zone, preheating zone, intense reacting zone, and product zone. It could be inferred that the preheating zone induced by the hot product in the cavity acts as a radical source and heat source for combustion stabilization in the cavity. In addition, the transient flame morphology analysis was carried out from the high-speed photography results. The flame base position and flame spreading angle were abstracted from the binarized image. The results showed that there existed a subtle flame flashback phenomenon for the flame base, whereas the flame spreading angle maintained a relatively constant value with a small deviation. The frequency-domain analysis revealed that both the flame base position and flame spreading angle reached the peak amplitude in the direct current component (0 Hz), exhibiting no explicit high-amplitude resonance component within 1500 Hz compared to those in the diffusion combustion mode. In all, the supersonic premixed flame presented a stabilized flame morphology.

Corner Separation Control Using Sweeping Jet Actuator in a Compressor Cascade

Abstract In the pursuit of advancing active flow control (AFC) technology, a more promising nonsteady actuator known as sweeping jet actuator (SJA) has been equipped on the endwall to manage the corner separation in a compressor cascade, while the steady jet actuators with holes were used for comparison. Experimental investigations have been conducted in a low-speed wind tunnel with an inflow Mach number of 0.23. Five-hole probe measurements and the oil flow visualizations were carried out to demonstrate the performance and physics of steady and unsteady blowing on controlling the corner separation. The transient data at the measurement plane were also obtained using the dynamic pressure sensors to get insight into the unsteady characteristics. Variation of jet momentum of both actuators and the excitation frequency of the SJA allows determination of favorable control parameters. Using the SJA with only 0.13% of the cascade mass flow, the total pressure loss which takes the additional energy input into account is reduced by 14.2%, while the steady jet achieves a reduction of 8.4%. The results highlight the superior characteristics of the SJA in controlling the corner separation.

Direct Numerical Simulation Database of High-Speed Flow over Parameterized Curved Walls

This study presents a direct numerical simulation (DNS) database of high-speed turbulent boundary layers (TBLs) subject to pressure gradients due to parametrically varied backward-facing and forward-facing wall curvatures, with an inflow Mach number of 4.9 and a friction Reynolds number of Reτ≃1000 immediately before the onset of wall curvature. The Mach and Reynolds numbers are significantly higher than those reported in the literature for the DNS of pressure-gradient TBLs. The flow conditions and baseline wall geometries are representative of experiments in the high-speed blowdown wind tunnel at the National Aerothermochemistry Laboratory at Texas A&M University. The wall steepness of the baseline geometry for both the backward-facing and forward-facing walls was systematically varied to cause attached, incipiently separated, and fully separated flows. Precomputed flow statistics, including turbulent kinetic energy budgets, are available on the website of the Turbulence Modeling Resource of the NASA Langley Research Center, allowing other investigators to query any property of interest.

On the airfoil leading-edge noise reduction using poro-wavy leading edges

This paper presents numerical studies on airfoil leading-edge turbulence interaction noise reduction using poro-wavy leading edges. Three different bionic treatments including wavy leading edges, porous leading edges, and a novel combined poro-wavy leading edges are modeled. The turbulent flow field is solved using the improved delayed detached eddy simulation method. The aerodynamic noise is predicted using the Ffowcs Williams and Hawkings acoustic analogy theory. The inflow Mach number is approximately 0.12 with an angle of attack of 0°, and the chord-based Reynolds number is 400 000. The present numerical method is first validated against experimental data and previous studies. Then the effects of the three bionic treatments on the aerodynamic performance and the aeroacoustic performance are analyzed. The results show that all the three bionic treatments will increase the mean drag of the airfoil, especially for the airfoils with porous treatment, while the lift and drag fluctuations are significantly reduced by the three bionic treatments. The wavy leading edges are found to be more effective for the reduction of broadband noise, while the porous leading edges are more effective for the reduction of the tonal noise. For the poro-wavy leading edges, both the tonal noise and broadband noise are significantly reduced, which means that the combined poro-wavy leading edges possess both the advantages of the wavy and porous treatments. The underlying flow mechanisms responsible for the noise reduction are finally analyzed in detail.

Combustion Characteristics and Application Performance of JP-10/Nano-Sized Aluminum Gel-Fueled Scramjet

Using nano-sized energetic materials as fuel additives to hydrocarbon gel fuel has gained in popularity due to the advantages of higher energy density and shorter ignition delay times. Combustion characteristics and application performance of gel fuel containing 16% wt aluminum nanoparticles were investigated on a direct-connect scramjet platform with an inflow Mach number of 2, an inflow total temperature of 1700 K, and a fuel equivalence ratio ranging from 0.6 to 1.0. The gel fuel could be stably injected and atomized efficiently by a nitrogen-assisted injector. The combustion efficiency of JP-10+Al gel fuel was 4.27% higher than that of JP-10 gel fuel, while the addition of aluminum increased the density specific impulse by 6.34%, indicating that the addition of aluminum nanoparticles indeed presented a positive effect on combustion and engine performance. The wall deposition phenomenon of gel fuel was particularly slight as compared to that of slurry fuel cases. Unlike nanoaluminum-blended slurry fuel, nanoaluminum-blended gel fuel controlled the potential risk of increasing thermal load to the least degree within the scope of existing research, with wall heat flux increasing by 7.43% solely as a result of elevated airflow temperature, indicating its advantageous application prospect.

Effects of inflow Mach numbers on shock train dynamics and turbulence features in a backpressured supersonic channel flow

Investigations on shock train dynamics and relevant turbulence features in a backpressured supersonic channel flow are carried out using direct numerical simulation for three inflow Mach numbers of Ma0= 1.61, 2.0, and 2.45. As Ma0 increases, the shock train undergoes a structural change characterized by the leading shock which changes from the symmetric “λ” (Ma0=1.61) to the symmetric “X” (Ma0=2.00) and then to the asymmetric “X” pattern (Ma0=2.45). The symmetry breaking of shock structures induces asymmetric separation, which significantly alters the distribution characteristics of wall variables such as wall pressure and friction. To examine the unsteady behaviors of the shock train, a mode decomposition technique, namely, reduced-order variational mode decomposition [Liao et al., J. Fluid Mech. 966, A7 (2023)], is adopted taking its merit of adaptively extracting time-frequency features of dynamic systems. The modal analysis reveals that the shock train system exhibits significant centralization of low-frequency energy. Specifically, two basic types of low-frequency oscillation modes dominate the unsteady motion of the shock train: one depicts overall translating oscillation while another represents accordion-like oscillation. The analysis of turbulent kinetic energy shows that turbulence amplification is mainly dominated by the interaction of the decelerating mean flow with streamwise velocity fluctuations in the vicinity of the leading shock for all three cases, which is unaffected by the symmetry breaking of shock structures.

Effect of Leading-Edge Erosion on the Performance of Transonic Compressor Blades

In this paper, an experimental and numerical investigation of the effect of leading-edge erosion in transonic blades was performed. The measurements were carried out on a linear blade cascade in the Transonic Cascade Wind Tunnel of DLR in Cologne at two operating points with an inflow Mach number of 1.05 and 1.12. The numerical simulations were performed by ANSYS Germany. The type and specifications of the erosion for the study were derived from real engine blades and applied to the leading edges of the experimental cascade blades using a waterjet process, as well as modeled in detail and meshed within the numerical setup. Numerical simulations and extensive wake measurements were carried out on the cascades to evaluate the aerodynamic performance. The increase in losses was quantified to be 4 percent, and a reduction in deflection and a rise in pressure were detected at both operating points.

Direct aeroacoustic simulation of a flow through an expanding pipe with orifice plates

Expanding pipes with orifice plates are utilized as silencers for fluid machinery. However, tonal sounds can be generated from flows through the silencers with such a configuration. To clarify the mechanism of tonal sound generation from a flow over a circular expanding pipe with orifice plates, flow and acoustic fields were directly solved on the basis of the compressible Navier-Stokes equations. Prediction results demonstrated that tonal sounds occurred at multiple frequencies. The phase-averaged flow fields for each frequency indicated that vortex rings or spiral vortices were periodically shed in the flow around the orifice plates. Acoustic radiation occurred owing to the collision of these vortices with the orifice plates, which led to the occurrence of acoustic resonance in the expanding pipe. The predicted circumferential phase distributions of pressure fluctuations demonstrated the coexistence of different acoustic modes in the expanding pipe. The relationship between acoustic modes and vortical behaviors was presented. The effects of inflow Mach number on the vortex structures and acoustic modes for the resonance were also discussed.

Investigation of a Near-Field Cylinder Wake in the Subsonic, Transonic, and Supersonic Regimes

The near wake of a circular cylinder in a high-Reynolds-number regime is investigated using the particle image velocimetry technique for an inflow Mach number spanning 0.3 to 2. The mean flow and turbulent kinetic energy field for different Mach number cases are presented. Furthermore, the influence of the Reynolds and Mach number effects is examined based on the mean separation point along the cylinder surface, which is extracted from complementary schlieren visualization measurements, and the literature. The wake width for the supersonic inflow case is found to be an order of magnitude smaller physically than the subsonic case with an identical cylinder diameter. Additionally, the swirl strength criterion is used to identify the eddies present within the wake, and the eddy size and convection velocity are characterized based on the spatiotemporal correlations. The result suggests that the mean wake features for the subsonic and supersonic cases are very different. Moreover, the interaction of eddies generated at the cylinder’s top and bottom shear layers leads to the large-scale features present in the wake, and the differing paths followed by the eddies between the subsonic and supersonic cases are responsible for the variation in the observed wake width.

Effect of wavy leading edges on airfoil tonal noise

This work presents numerical studies on the effect of wavy leading edges on airfoil instability tonal noise. Large eddy simulations are conducted for NACA0012 (National Advisory Committee for Aeronautics) airfoils with straight and wavy leading edges. The far-field aerodynamic noise is predicted using the Ffowcs Williams and Hawkings acoustic analogy theory. The inflow Mach number is approximately 0.17 with an angle of attack of 4.3°, and the chord based Reynolds number is 600 000. The present numerical method is first validated by existing numerical results and a semi-empirical model for the straight baseline airfoil. Then, the effects of wavy leading edges on the aerodynamic performance, aeroacoustic performance, and noise reduction mechanisms are discussed in detail. The wavy leading edge is found to be detrimental to the mean aerodynamic performance with a decreased lift and increased drag. However, the lift fluctuations are significantly reduced. The instability tonal noise and its harmonic are totally removed by the wavy leading edges, while an extra broadband hump appears at the middle frequency. The low frequency and high frequency noise are also increased by the wavy airfoil. The laminar separation bubble is removed, and the flow separation is decreased by the wavy airfoil in the vicinity of the trailing edge. The pressure fluctuations around the trailing-edge are considerably diminished, and the spanwise coherence spectra are also significantly reduced by the wavy airfoil. Furthermore, it is noteworthy that the coherence reduction spectrum at the source is almost consistent with the far-field noise reduction spectrum.

Optimal two-dimensional roughness for transition delay in high-speed boundary layer

The influence of surface roughness on transition to turbulence in a Mach 4.5 boundary layer is studied using direct numerical simulations. Transition is initiated by the nonlinearly most dangerous inflow disturbance, which causes the earliest possible breakdown on a flat plate for the prescribed inflow energy and Mach number. This disturbance primarily comprises two normal second-mode instability waves and an oblique first mode. When localized roughness is introduced, its shape and location relative to the synchronization points of the inflow waves are confirmed to have a clear impact on the amplification of the second-mode instabilities. The change in modal amplification coincides with the change in the height of the near-wall region where the instability wave speed is supersonic relative to the mean flow; the net effect of a protruding roughness is destabilizing when placed upstream of the synchronization point and stabilizing when placed downstream. Assessment of the effect of the roughness location is followed by an optimization of the roughness height, abruptness and width with the objective of achieving maximum transition delay. The optimization is performed using an ensemble-variational (EnVar) approach, while the location of the roughness is fixed upstream of the synchronization points of the two second-mode waves. The optimal roughness disrupts the phase of the near-wall pressure waves, suppresses the amplification of the primary instability waves and mitigates the nonlinear interactions that lead to breakdown to turbulence. The outcome is a sustained non-turbulent flow throughout the computational domain.

Thermochemical non-equilibrium flow characteristics of high Mach number inlet in a wide operation range

The high-temperature non-equilibrium effect is a novel and significant issue in the flows over a high Mach number (above Mach 8) air-breathing vehicle. Thus, this study attempts to investigate the high-temperature non-equilibrium flows of a curved compression two-dimensional scramjet inlet at Mach 8 to 12 utilizing the two-dimensional non-equilibrium RANS calculations. Notably, the thermochemical non-equilibrium gas model can predict the actual high-temperature flows, and the numerical results of the other four thermochemical gas models are only used for comparative analysis. Firstly, the thermochemical non-equilibrium flow fields and work performance of the inlet at Mach 8 to 12 are analyzed. Then, the influences of high-temperature non-equilibrium effects on the starting characteristics of the inlet are investigated. The results reveal that a large separation bubble caused by the cowl shock/lower wall boundary layer interaction appears upstream of the shoulder, at Mach 8. The separation zone size is smaller, and its location is closer to the downstream area while the thermal process changes from frozen to non-equilibrium and then to equilibrium. With the increase of inflow Mach number, the thermochemical non-equilibrium effects in the whole inlet flow field gradually strengthen, so their influences on the overall work performance of the high Mach number inlet are more obvious. The vibrational relaxation or thermal non-equilibrium effects can yield more visible influences on the inlet performance than the chemical non-equilibrium reactions. The inlet in the thermochemical non-equilibrium flow can restart more easily than that in the thermochemical frozen flow. This work should provide a basis for the design and starting ability prediction of the high Mach number inlet in the wide operation range.

Numerical investigation on the initiation of oblique detonation waves in stoichiometric methane–air mixtures

The study investigates the oblique detonation characteristics in a stoichiometric methane–air mixture employing reactive Euler equations coupled with a detailed chemical reaction model. The analysis evaluates the initiation variation for both the inflow Mach number and the wedge angle. The results show that an increase in the Mach number from M0 = 9 to M0 = 12 changes the oblique shock-to-detonation transition from an abrupt to a smooth type. Furthermore, the distinction between high and low Mach numbers is explored through a comparison of the induction length along the wave surface. Additionally, the transition is affected by changes in wedge angle elevation under high Mach numbers, while a large wedge angle generates a second detonation wave surface due to a strong reflection shock downstream under low Mach numbers. Finally, the initiation length is quantitatively compared across varying inflow Mach numbers and wedge angles, which can benefit future investigations of a methane-fueled oblique detonation engine.

Comparison of LDA and hot-wire measurements at moderate subsonic mach numbers

A detailed comparison of Hot-Wire-Anemometry (HWA) and Laser-Doppler-Anemometry (LDA) measurements in the wake of an airfoil is presented. An evaluation of the capabilities of both techniques in measuring three-dimensional turbulence in turbomachinery applications is made and potential application pitfalls are highlighted. It is shown that the HWA tends to underestimate the turbulence intensity due to damping effects, while the LDA tends to overestimate the turbulence intensity due to the optical setup. The wake flow of a single airfoil at inflow Mach numbers of 0.35 and 0.45 and corresponding Reynolds numbers of 480,000 and 630,000 are used as a reference case. Different HWA probe head designs and various wire diameter are considered, as well as different setups for the LDA measurement device. In a comparison of the two techniques, the turbulence intensity measured differs by up to <inline-formula><mml:math xmlns:mml="http://www.w3.org/1998/Math/MathML" display="inline" overflow="scroll"><mml:mn>1.3</mml:mn></mml:math></inline-formula> percentage points (pp) in the freestream region and 4.0 pp within the wake. This is primarily due to the resolution of high-frequency fluctuations. For both measuring techniques, potential sources of errors are highlighted, especially regarding the application to turbomachinery flows.

Analysis of muzzle flow field of asynchronous firing of multi barrel aircraft gun under strong convection

Based on the N - S equation and the k-ε model, this article uses the dynamic grid technology to finish the numerical simulation of projectile-barrel coupling for muzzle flow field of the double-barreled aircraft gun during asynchronous firing under different incoming Mach numbers and different firing intervals. The influence law of muzzle flow field on projectile motion is studied. The results show that when there is supersonic incoming flow, the radial force on the projectile will increase with the increase of incoming Mach number. At the same time, the larger the Mach number, the earlier the radial force of the projectile will be affected. Under the same inflow Mach number, the radial force on the projectile when the double-barreled aircraft gun is fired asynchronously is significantly lower than that when fired synchronously. The larger the firing interval, the smaller the radial force on the projectile.

Effects of Inlet Incidence Perturbations on Compressor Cascade Performance using Adaptive Sparse Grid Collocation

The effects of inflow variations due to the working environment and flight attitude changes on turbomachines are considerable in the real world. Nevertheless, uncertainty quantification can be adopted to assess mean performance changes and perform the aerodynamic shape design as well as optimization. Thus, an uncertainty quantification method of adaptive sparse grid collocation (ASGC) was first introduced to address the inflow uncertainties’ effect issue effectively and accurately. Then, ASGC was utilized to evaluate the impacts of inlet incidence perturbations at different perturbation scales and reference inflow Mach numbers on the aerodynamic performance of a controlled diffusion cascade. The results showed that compared with the Monte Carlo simulation and static sparse gird collocation, the statistical accuracy and response accuracy of ASGC were maintained, and meanwhile its model construction efficiency was significantly improved because of the nested adaptive sampling feature. Under the perturbations of inlet incidences with high reference incidences, the mean aerodynamic loss always aggravates. The changes in aerodynamic loss nonlinearly depend on the inlet incidence perturbations, and the nonlinear dependence becomes greater when the perturbation scale. expands. At the same perturbation scale, the nonlinear dependence on the inlet incidence perturbations is further enhanced when the reference inflow Mach number rises. Finally, uncertainty quantification of the flow field revealed that the fluctuation of flow accelerations at the leading edge plays a fundamental role in determining the uncertainty of the aerodynamic loss.