Shock Boundary Layer Interaction Research Articles

Experimental Evaluation of Isolated Intake for Hypersonic Air- Breathing Vehicle

Isolated intake of an air-breathing conceptual hypersonic demonstrator vehicle has been experimentally characterized simulating the entry Mach number and the entry flow direction. A 1:8 scale intake model having provision to vary the entry area of the internal compression portion by means of actuating the intake cowl has been tested in the supersonic freestream Mach numbers 3, 3.44 and 3.9. Tests were carried out in the 0.3 m x 0.3 m trisonic wind tunnel at National Aerospace Laboratories, Bangalore. The result pertaining to M¥ = 3.44 is discussed in this paper. The effect of cowl position on the intake starting characteristics and performance are analyzed using wall pressure measurements, schlieren and oil flow visualization. Intake unstart occurred before the cowl reaching to its fully open position with the original hinge location. Essentially the strength of cowl leading edge shock increases as the cowl opening angle increases. The strong shock-boundary layer interaction in the convergent portion of intake aids the growth of the separation bubble in the flow path and that cause the intake to unstart. Improvement of intake starting by means of shifting the cowl hinge location downstream was attempted. Shifting the cowl hinge location downstream has improved the intake starting characteristics. On the improved configuration, the effect of combustion induced back pressure and the effect of cowl opening speed on intake unstart have been analyzed and quantified.

Shock Induced Flow-Separation in Hypersonic Intakes at Off-Design Conditions

High-speed turbulent flow in hypersonic intakes encounters flow-separation due to shockboundary layer interaction. Presence of separation bubble changes the pressure distribution over the surface which can lead to deterioration of aerodynamic performance of intake and thus jeopardize the stability of vehicle. In this study, we employ the physics-based shock-unsteadiness k-e model to compute separation bubble size in a mixed compression two-ramp intake. The design point Mach number of 6 at 30 km altitude with zero angle of attack is considered. Changes in separation bubble length is analyzed at different off-design conditions by varying the flight Mach number, angle of attack and cruise altitude. The changes observed in the separation length by variation of above flight parameters is studied in terms of the physically relevant canonical parameters of upstream Mach number, Reynolds number Red and the ratio of wall to recovery temperature. CFD simulations are performed for the entire range of canonical and flight parameters. Wall-to-recovery-temperature ratio is found to be an important governing parameter for separation length studies. The largest shock-induced separation is observed for the case of off-design flight Mach number of 8. The corresponding effect on the lift and drag generated by the hypersonic intake are quantified.

Deep learning reconstruction of pressure fluctuations in supersonic shock–boundary layer interaction

The long short-term memory deep-learning model is applied to supersonic shock–boundary layer interaction flow. The study aims to show how near-wall pressure fluctuations can be reconstructed from reduced (under-sampled) datasets of pressure signals. Predicting pressure fluctuations from reduced datasets could allow predictions using less expensive simulations and experiments. The training of the deep learning model is based on direct numerical simulations of supersonic ramp flows, focusing on the regions upstream of and around the shock–boundary layer interaction region. During the pre-processing stage, cubic spline functions increase the fidelity of the sparse signals and feed them to the long-short memory model for an accurate reconstruction. Comparisons are also carried out for different sparsity factors and assess the model's accuracy both qualitatively through the pressure signals and quantitatively using the root mean square error and the power spectra. The deep learning predictions are promising and can be extended to include other aerodynamic or aeroelastic parameters of interest.

Research on time series prediction of the flow field in supersonic combustor based on deep learning

The detection of supersonic combustion flow field provides valuable reference information for the efficient combustion of hypersonic vehicles. From the perspectives of scramjet operation during flight as well as experiment where the measured information is limited, flow field prediction using deep learning technology is a promising method that can provide future flow field evolution in scramjet combustors. This study proposes a convolutional neural network model of multi-temporal pressure information fusion to predict combustion flow field based on wall pressure. In addition, it analyzes the effect of various wall pressures on the flow field prediction. An experimental dataset of supersonic ground-pulse wind tunnel tests with different injection pressures is built through the combustion chamber wall under self-ignition conditions of a hydrogen-fueled scramjet. The results showed that the predicted flow field of the combustor agrees with the ground test, average correlation coefficient reached more than 90% and when the wall pressure number is 32, the average peak signal to noise ratio reaches about 23. and the main wave structure and turbulent boundary layer region are restored even if there is a shock boundary layer interaction.

An Insight into Quasi-Two-Dimensional Flow Features Over Turbine Blading From the Works of Jonathan Paul Gostelow

Abstract The flow through the predominantly two-dimensional geometries of cascades of blades is intrinsically three-dimensional and unsteady. Direct Numerical Simulation, Large Eddy Simulations, and time-resolved Particle Image Velocimetry provide access to the full flow physics, relevant to aerodynamic loss and heat management. Such studies build upon earlier insight drawn from quasi-two-dimensional investigations that identified the key areas where progress in understanding was most needed. These areas stretch across the full passage, from the leading edge of the blade to the passage outflow. Streamwise surface vorticity, transition, the calmed region, shock–boundary layer interaction, and vortex shedding are considered in detail, specifically (i) on what gaps in their physical understanding the works of Jonathan Paul Gostelow exposed and (ii) what gaps were present in the two-dimensional computational approaches used to represent these flows in these works. These useful insights are obtained from the geometrically simpler settings of circular cylinders in cross-flow and from flat plate experiments, as well as from cascades of blades. This paper presents an overview of the physical understanding of the flow features that underpins the more recent time-resolved three-dimensional investigations, led by the late Emeritus Professor Jonathan Paul Gostelow. This work celebrates some of Paul Gostelow’s 50 + years of turbomachinery research achievements and develops awareness about their significance toward reaching a more complete knowledge of the flow physics in turbomachinery, using the more recent time-resolved three-dimensional modeling capability of Computational Fluid Dynamics.

Numerical study of transverse hydrogen injection in high-speed reacting crossflow

A high-speed compressible solver capable of solving detailed chemical reaction mechanisms is developed by coupling the open-source computational fluid dynamic toolbox OpenFOAM® and Cantera 2.5.0. A sonic hydrogen jet discharging from a circular injector into a high enthalpy supersonic crossflow over a flat plate is selected as a test case for the developed solver. The incoming boundary layer is laminar, and an adverse pressure gradient-induced transition is expected due to transverse injection. The test case is selected to serve two purposes. First, to validate the developed solver. Second, to inspect the capability of Reynolds-Averaged Navier–Stokes (RANS) in predicting the flame characteristics in high-speed flows involving laminar to turbulent transition. The present study features three-dimensional RANS simulations with Shear Stress Transport (SST) k–ω and Langtry–Menter SST k–ω turbulence models, with three values of inlet turbulent intensity: I = 0.5, 1, and 2. Analysis showed that laminar to turbulent transition plays a significant role in the resulting flame structure. A fully turbulent SST k–ω model showed several discrepancies from the experiment, especially near the boundary layer. On the other hand, the Langtry–Menter SST k–ω model predicts transition onset and hence the flame structures accurately. Furthermore, the transition onset and the flame structure strongly depend on I. The low-velocity recirculation regions near the injector aid in flame stabilization upstream of the injector. At the same time, the horseshoe vortex dictates the flame spread in a spanwise direction. The reflected shock–boundary layer interaction helps in flame stabilization downstream of the injector.

Grey area in embedded wall-modelled LES on a transonic nacelle-aircraft configuration

A scale-resolving hybrid RANS–LES technique is applied to an aircraft-nacelle configuration under transonic flow conditions using the unstructured, compressible TAU solver. In this regard, a wall-modelled LES methodology is locally applied to the nacelle lower surface to examine shock-induced separation. To circumvent the grey-area issue of delayed turbulence onset, a Synthetic Turbulence Generator (STG) is used at the RANS–LES interface. Prior to the actual examinations, fundamental features of the simulation technique are validated by simulations of decaying isotropic turbulence as well as flat plate flows. For the aircraft-nacelle configuration at a Reynolds number of 3.3 million, a sophisticated mesh with 420 million points was designed which refines 32 % of the outer casing surface of the nacelle. The results show a development of a well-resolved turbulent boundary layer with a broad spectrum of turbulent scales which demonstrates the applicability of the mesh and method for aircraft configurations. Furthermore, the necessity of a low-dissipation low-dispersion scheme is demonstrated. However, a noticeable drop of the surface skin friction downstream of the STG motivates further research on the impact of the interface modelling on the shock–boundary layer interaction.

Transport-Equation-Based Shock Sensor for Shock–Turbulent Boundary-Layer Interaction

High-speed flows are often associated with shock–boundary-layer interaction (SBLI), which causes excessive vehicle surface heating and total pressure loss. Therefore, precise computational prediction of shock interactions becomes essential. Commonly used simulation techniques, including numerical schemes and turbulence models, give erroneous results at shock discontinuities. Local modifications are applied at shocks based on their location and strength. Most shock sensors in the literature predict shock location but do not give shock properties. We propose a shock function to locate shocks and estimate their strength in terms of density ratio by solving a transport equation, along with the governing equations of fluid flow. The sensitivity of the shock function to the transport equation’s parameters and grid resolution are explored. The shock function is used to define a shock sensor that can identify thin and sharp shocks. The shock function and shock sensor are then applied to an oblique shock impinging on a turbulent boundary layer. We use the shock unsteadiness modified model, which uses the estimated shock strength to predict the turbulence levels at shocks accurately. We study the effect of nonadiabatic wall temperature on the SBLI region, and the results are compared with available Direct Numerical Simulation (DNS) data.

Leading-edge bluntness effects on the hypersonic flow over the double wedge at multiple aft-wedge angles

The present numerical investigation focuses on the leading-edge bluntness effects on the double wedge with varied aft-wedge angles exposed to low enthalpy hypersonic free stream conditions. The bluntness ratio in this study varies, ranging from R/L1 = 0 (sharp leading edge) to R/L1 = 0.577 (maximum allowable bluntness), along with the aft-wedge angle varying between θ2 = 45° and 60°. Noticeably, even a small bluntness ratio can completely change the shock interaction pattern compared to its sharp geometrical counterpart due to a detached leading-edge shock, enlarged separation bubble, and location of various shock waves concerning it. Critical bluntness ratios exist for the low aft-wedge θ2 = 45° angle, but increasing the aft-wedge angle makes the flow field highly unsteady for some bluntness ratios. Nevertheless, these bluntness ratios for such double-wedge configurations are reported using the mean of separation bubble size. Moreover, this work unravels the cause of such unsteadiness for the unsteady flow fields using the spatial-temporal evolution of the wall pressure distribution and fast Fourier transform of the pressure fluctuation signal at the compression corner and supports the deduced observation with the help of energy-based proper orthogonal decomposition. The increased shock–boundary layer interaction strength moves the separation point upstream beyond the junction of cylindrical bluntness and inclined fore-wedge surface, accompanying sudden change in its direction of motion that perturbs the shear layer that set to a self-sustained, highly unsteady flow field.

Assessment of the shear stress transport dynamic ℓ2−ω delayed detached eddy simulation in Bachalo–Johnson flow with shock-induced separation

Driven by the need for simulating compressible flows, Germano identity-based [Z. Yin and P. A. Durbin, “An adaptive DES model that allows wall-resolved eddy simulation,” Int. J. Heat Fluid Flow 62, 499–509 (2016)] and Vreman operator-based [Bader et al., “A hybrid model for turbulence and transition, with a locally varying coefficient,” Flow, Turbul. Combust. 108, 935–954 (2022)] dynamic ℓ2−ω delayed detached eddy simulation (DDES) formulations are constructed on the k−ω shear stress transport (SST) model. The Bachalo–Johnson transonic axisymmetric bump is simulated to assess the models’ capability in handling the compressible boundary layers under pressure gradient and transonic shock–boundary layer interaction. The new dynamic ℓ2−ω DDES formulation based on k−ω SST overcomes the issues of freestream sensitivity and inaccurate compressible boundary layer profile observed in the original k−ω (88) based model. The new SST-based dynamic model using the Vreman operator to compute the model coefficient (Vreman-dynamic model) has superior performance against Germano identity-based model due to its capability of suppressing the subgrid viscosity during the initial development of a separating shear layer. The Vreman-dynamic model predicts a reattachment location similar to the zonal improved-DDES/direct numerical simulation approach by Spalart et al. [“Large-eddy and direct numerical simulations of the bachalo-johnson flow with shock-induced separation,” Flow, Turbul. Combust. 99, 865–885 (2017)] on a much coarser mesh demonstrating its potential for application in industrial flows.

Vibration of a thin panel exposed to ramp-induced shock-boundary layer interaction at Mach 2

Flight vehicles that operate in the supersonic regime can be subject to adverse fluid–structure interactions due to their lightweight design. The presence of geometric obstructions such as control surfaces or fins can induce shocks that can interact with the boundary layer, leading to flow separation. This work investigates experimentally the interaction between a compliant panel in a Mach 2 flow under a compression ramp-induced shock-wave/boundary-layer interaction (SBLI). Thin brass panels of different thickness are investigated in a wind tunnel. Tests are performed both with and without a 20° compression ramp installed. This direct comparison allows characterization of the effect of the SBLI on the system dynamics. High-speed stereoscopic digital image correlation (DIC) and fast-response pressure sensitive paint (PSP) are used to obtain simultaneous measurements of full field deformation and surface pressure of the panels. The panel vibration is dominated by the first bending mode. Despite the forcing of the separation shock foot, the presence of the SBLI does not significantly modify the operational deflection shape, frequency, and amplitude of the dominant vibration mode, beyond what is observed for the no-SBLI case. On the other hand, analysis of the shock foot motion shows that the shock primarily oscillates at the first natural frequency of the panel. This leads to the conclusion that the shock foot oscillation is driven by the panel vibration in a one-way coupling mechanism. The SBLI does modify the higher modes, which is likely due to localized forcing by the separation shock foot. Full-field surface pressure predictions are made using first order piston theory. Results show that the fluid–structure interaction is dominated by the large region of attached flow upstream of the shock foot.

Flow Characteristics and Parameter Influence of the Under-Expansion Jet on Circulation Control Airfoil

The enhancement in the jet pressure ratio and jet velocity contributes to expanding the control efficiency and control boundary of circulation control airfoil under high subsonic incoming flow. However, because of an excessive jet pressure ratio, the jet separates prematurely on the Coanda surface, resulting in control failure. In a bid to improve the adhesion capability of the jet under a high pressure ratio, a circulation control airfoil with a converging nozzle and back-facing step structure at the trailing edge was numerically simulated based on the Reynolds averaged Navier−Stokes equation (RANS), and a study was conducted on the complex flow structure of the under-expansion jet on the Coanda surface and the impact of design parameters such as jet pressure ratio, ellipticity, and nozzle height on the jet separation. The results show that the back-facing step provides an expansion space for the under-expansion jet and changes the shock-boundary layer interaction form. As the jet pressure ratio and nozzle height increase, the size of the shock cell increases, the strength of the intercepting shocks on both sides increases, and Mach reflection occurs, resulting in jet stratification and in a decline in the adhesion capability of the jet. The combination design of proper ellipticity and the back-facing step contributes to forming a closed low-pressure vortex area behind the step and promote jet attachment. Reducing the nozzle height can improve the adhesion capability of the jet under a high pressure ratio.

Fluid structure interactions generated by an oblique shock impinging on a thin elastic panel

The flow structure coupling generated by an impinging shock boundary layer interactions (SBLI) flowfield over a rectangular panel is investigated. The impinging oblique shock is generated by an 8° shock generator placed incident to a Mach 2.5 flow. The shock strength is large enough to generate a mean separation that is nearly two-dimensional over the panel span. The panel is configured such that the panel oscillation amplitude is much smaller than the panel and incoming boundary layer thicknesses. This resulted in a weak coupling between the flowfield and structural response, and forms a relatively simpler configuration to study. Two multivariate measurement campaigns are performed to capture the mean and dynamic flowfield and panel response. The first campaign simultaneously measures the 2D panel surface pressure field, panel center-span deflection, and off-body velocity field at 10 Hz. The second campaign measures simultaneously the 2D panel surface pressure field and panel strain at two mid-span locations at 10 kHz. The acquisition rate is sufficient to resolve at least the first seven panel elastic models. These measurements provide comprehensive information about the fluid structure interaction (FSI) phenomenon from both aerodynamics and structural dynamics perspectives. Summarily, with the weak flow/structure interactions implemented in this study, there is no change in the mean separation size. However, the pressure fluctuations beneath the SBLI unit exhibit discrete elevations in the vicinity of different panel resonance mode frequencies. The greatest elevation pressure PSD occurs at the panel modes that overlaps with the separation bubble pulsation frequency band. Detailed coherence maps at and away from panel elastic modes and conditional statistics reveal critical insights into the mechanisms that cause the strengthening of the pressure fluctuations due to flow/structure coupling. It was found that the acoustic forcing by the cavity resonance has an important contribution to the dynamics of the surface pressure and panel oscillations.

Investigations of shock–boundary layer interaction dynamics using high-bandwidth pressure field imaging

The large-scale pulsations of shock-induced separation with length scale that significantly exceeds the incoming boundary layer thickness are investigated. The shock–boundary layer interaction (SBLI) unit is generated by an inward-turning axisymmetric compression ramp at an inflow Mach number of 2.5. A substantial region surrounding the centre azimuth exhibited mean and dynamic flow features that are consistent with two-dimensional separation. Two-dimensional highly resolved maps of surface pressure field are obtained using fast-response pressure-sensitive paint fluorescence imaging at 40 kHz repetition rate. The measurement domain covered significant regions of the incoming boundary layer through the relaxing boundary layer downstream of the reattachment as well as over 25 boundary layer thicknesses in the azimuthal direction. These measurements provide new insights into the spanwise coupling of the SBLI unit in addition to its inherent dynamics. The power spectral density (PSD) of the centreline pressure exhibits very good agreement with theoretical models and complementary measurements using fast-response pressure transducers, which served to validate the pressure field measurements. Detailed examination of the PSD reveals strong agreement with the literature, which includes the peak Strouhal number of the separation and reattachment shock motions as well as the downward frequency shift along the separation bubble. Furthermore, the pressure fluctuation maps reveal streamwise-elongated structures just downstream of the ramp leading edge that persist well downstream of the reattachment. A time sequence of conditional average pressure fluctuation maps is constructed surrounding isolated pressure excursions in the intermittent region. This sequence, along with two-point cross correlation analysis, provides critical information about the flow processes that drive the separation bubble pulsations in the SBLI units with large separation scales. Overall, the imbalance in the mass within the separation bubble appears to be the critical mechanism that drives the separation bubble pulsations. Furthermore, the pressure perturbations originating at azimuthally offset locations are also observed to influence the separation bubble dynamics.

High order discontinuous Galerkin simulation of hypersonic shock-boundary layer interaction using subcell limiting approach

Numerical simulation of shock-boundary layer interactions in the hypersonic regime has for long been a challenge, plagued by high sensitivity of the surface predictions to grid and numerical details. In the course of developing adaptive high order methods to mitigate this issue, discontinuous Galerkin method has emerged as a promising alternative to the conventionally used finite volume method. However, till date, among the reported finite element method-based simulations of these flows, only a few have presented high order simulations using a higher than linear basis. In this work, a subcell limiting technique devised for the compressible Navier-Stokes equations is extended for use in the hypersonic regime. The proposed extension is tested by performing simulations of several shock-boundary layer interaction cases in the freestream Mach number range 6 to 14 using a quadratic basis. For a set of cases at Mach 6, grid converged simulations are presented that agree reasonably well with experimental data and other numerical solutions. More importantly, simulations of a Mach 11 case that are performed using two different flux schemes are shown to yield nearly identical predictions of the surface pressure and heat transfer, suggesting that high order solutions can reduce the sensitivity of numerical predictions to the computational details. Grid convergence studies of two cases at Mach numbers 11 and 14 led to observing large scale oscillations in the separation bubble in finer grid simulations, suggesting that explicit time integration may not always be suitable for high order computation of hypersonic separated flows at very high Mach numbers.

Turbulence modeling of 3D high-speed flows with upstream-informed corrections

Turbulence modeling has the potential to revolutionize high-speed vehicle design by serving as a co-equal partner to costly and challenging ground and flight testing. However, the fundamental assumptions that make turbulence modeling such an appealing alternative to its scale-resolved counterparts also degrade its accuracy for practical high-speed configurations, especially when fully 3D flows are considered. The current investigation develops a methodology to improve the performance of turbulence modeling for a complex Mach 8.3, 3D shock boundary layer interaction (SBLI) in a double fin geometry. A representative two-equation model, with low-Reynolds number terms, is used as a test-bed. Deficiencies in the baseline model are first elucidated using benchmark test cases involving a Mach~11.1 zero pressure gradient boundary layer and a Mach~6.17 flow over an axisymmetric compression corner. From among different possibilities, two coefficients are introduced to inhibit the non-physical over-amplification of (i) turbulence production and (ii) turbulence length-scale downstream of a shock wave. The coefficients rely on terms already present in the original model, which simplifies implementation and maintains computational costs. The values of the coefficients are predicated on the distribution of turbulence quantities upstream of the shock; this ensures that the modifications do not degrade the model predictions in simpler situations such as attached boundary layers, where they are unnecessary. The effects of the modifications are shown to result in significant improvements in surface pressure and wall heat flux for the 3D SBLI test case, which contains numerous features not observed in 2D situations, such as 3D separation, skewed boundary layers and centerline vortices. Considerations on the inflow values of turbulence variables and mesh resolution are provided.

Flow control effect of pulsed arc discharge plasma actuation on impinging shock wave/boundary layer interaction

The current study investigates the control effect of the pulsed arc discharge plasma on the impinging shock wave and boundary layer interaction (SWBLI) generated by a 14° wedge in a Mach 2.5 flow. The response characteristics of SWBLI on pulsed arc discharge actuation were illustrated, and the controlling mechanism of shock-induced flow separation under different plasma power settings was revealed. The results, which were well validated by the relative published experiment, showed that when setting the exciting power density ph as 1.0 × 1011 W/m3, the oblique shock wave obtained an obvious fluctuation, and the foot of the reattachment shock wave was partially removed. In addition, as the controlling gas bubble passed through the interaction region, the reverse flow zone was enlarged, and the separation shock wave was shifted upward. When ph was set to 4.8 × 109 W/m3, the flow separation induced by SWBLI was effectively suppressed and the size of the reverse flow zone was significantly reduced. Moreover, as the energy input was increased, the arc-induced blast wave (BW) velocity was obviously enhanced. Additionally, it is further found that the arc plasma energy deposition density in the discharge region was the determining factor for SWBLI control, even for a relatively small exciting energy input. Better drag reduction of the flow field would be achieved with the pulsed arc discharge plasma of higher power density, and a drag reduction rate of nearly 10.05% was obtained at ph = 1 × 1011 W/m3 control condition.

Low-frequency unsteadiness of recompression shock structures in the diffuser of supersonic ejectors

Supersonic ejectors are passive gasdynamic devices that compress a low-pressure fluid by utilizing the kinetic energy of a high-pressure fluid in a variable area duct. The ejector consists of a primary supersonic nozzle in a mixing duct where the secondary flow is entrained and mixed. The mixed flow can undergo a series of recompression shocks resulting in a subsonic flow in the diverging portion to aid pressure recovery. Recompression shocks usually lead to unsteady shock boundary layer interactions. The performance of the ejector is influenced by shear layers, shock and expansion waves, and their mutual interactions. While existing literature has extensively dealt with mixing of the primary and secondary flows, the unsteadiness in flow resulting from recompression shocks has been seldom investigated. Fluctuations in pressure due to the unsteadiness of the shock often lead structural fatigue issues. This paper reports a detailed investigation on low-frequency unsteadiness of recompression shock using high-speed schlieren images and dynamic pressure measurements. Modal analyses using proper orthogonal decomposition and dynamic mode decomposition techniques are used to determine the dominant spatial modes and associated frequencies. Multimodal frequencies ranging between 80 and 300 Hz are observed. These findings are further corroborated by Fourier and wavelet transformations of the experimentally measured wall static pressure signals. Subsequently, scaling parameter is established for the dominant frequencies based on flow velocities upstream of the shock and the distance between two consecutive shocks. This results in a unique scaling frequency of 4.58% ± 18%, for the recompression shock independent of operating conditions.

Boundary layer control for low Reynolds number fan rig testing

Ultra-high bypass ratio turbofans offer significant reductions in fuel and pollution due to their higher propulsive efficiency. Short intakes might lead to a stronger fan-intake interaction, which creates uncertainty in stability at off-design conditions. Due to the prohibitive cost of full-scale experimental testing, subscale testing in wind tunnels is used to understand this behaviour. The low Reynolds number of subscale models results in unrepresentative laminar shock-boundary layer interactions. The boundary layer state thus needs to be conditioned to better represent full-scale transonic fans. This paper proposes the use of an inexpensive and robust flow control method for the suction side of a fan blade. Design guidelines are given for the location and height of the discrete roughness elements used to control the boundary layer state. This paper also presents a rapid experimental validation methodology to ensure and de-risk the application of the boundary layer trip to 3D rig blades. The experimental methodology is applied to a generic aerofoil representative of a fan tip section. The experimental method proves that it is possible to reproduce boundary layers and pressure distributions of a full-scale fan blade on a 1/10 subscale model. The results obtained confirm that the boundary layer trip method successfully promotes transition at the location representative of full-scale blades, avoiding unrepresentative laminar shock wave boundary layer interactions. This highlights the importance of conditioning boundary layers in low Reynolds number fan rig testing.

Aerodynamic Optimization and Mechanism Investigation on Performance Improvements in a Transonic Compressor Cascade

In this paper, a transonic compressor cascade was optimized to improve its aerodynamic performance. A new blade parameterization method with 16 control variables was first proposed to fit the shapes of the suction and the pressure side, as well as the leading edge. Then, the Kriging surrogate-model-based genetic algorithm (GA) was used to optimize the performance of the transonic cascade. The optimization algorithm is effective in reducing the total pressure loss while extending the working range of the cascade. The results show that the total pressure loss coefficient could be reduced by 11% at the best airflow angle and the working range could be extended by 6.9% for the optimized cascade in two-dimensional simulations. Similar improvement results could also be obtained in the simulations of their linear cascade cases. Detailed analyses show that the relative maximum thickness positions of the optimized blades move forward by about 10% to the leading edge, and the radii of curvature of the front half of the suction and pressure surfaces increase, compared with the initial blade. This makes the front half of the optimized blades look more closely like a wedge. Consequently, the passage shock strength is reduced and the shock changes from the passage normal shock to oblique shock. The weakened shock strength leads to the disappearance of the flow separation caused by the shock boundary layer interaction on the suction surfaces of the optimized blades, and results in a narrowed wake width at the outlet section.