Transonic Conditions Research Articles

Approach for Aerodynamic Gust Load Alleviation by Means of Spanwise-Segmented Flaps

Active gust load alleviation techniques exhibit a high potential in significantly reducing the transient gust loads on aircraft. In this work the aerodynamic potential of trailing-edge flaps and leading-edge flaps is numerically studied with the purpose to significantly reduce the structural gust loads. The utilized spanwise-segmented flaps represent slight modifications of existing devices for high-lift and maneuvering. The investigations based on unsteady Reynolds-averaged Navier–Stokes computations are conducted by employing a generic wing–fuselage aircraft configuration at transonic flow conditions. Idealized discrete “”-type vertical gusts that are relevant for the certification process are used as representative atmospheric disturbances. The focus of this paper is to introduce a practicable prediction method for required trailing- and leading-edge flap deflections for a significant mitigation of gust-induced wing loads. The three-dimensional flap deflections are determined by parametric two-dimensional simulations at representative wing sections. Different extensions of the estimation approach are investigated to assess the influence of the wing planform, the finite wing span, the aerodynamic phase lags, and the flap scheduling. It is shown that the trailing- and leading-edge flaps are promising in terms of alleviation of gust-induced wing bending and wing torsional moments, respectively. However, at high leading-edge flap deflections that are necessary for a full compensation of the wing torsional moment large-scale flow separation is identified. The introduced gust load alleviation approach indicates a good transferability between two-dimensional airfoil and three-dimensional wing aerodynamics for unsteady flap deflections.

Pressure-sensitive paint measurements on the cavity with passive control devices under transonic flow

Abstract In this study, the pressure-sensitive paint (PSP) technique, specifically the mesoporous-particle-based PSP, was employed to compare rectangular cavities with varying length-to-depth ratios (L/D) and different trailing edge shapes under transonic conditions. By utilizing PSP, comprehensive and quantitative pressure data were obtained, enabling the simultaneous observation of surface flow field distribution. The results obtained using PSP were found to be consistent with those obtained from conventional pressure sensors. The study revealed that the pressure distribution within the cavities changed with increasing L/D, and cavities with different trailing edge shapes demonstrated a reduction in pressure at the bottom region. Furthermore, the comparison of results obtained through the oil flow method corroborated the PSP findings, indicating that a beveled or sawtooth-shaped trailing edge of the cavity induced air flow deflection, effectively disrupting the upstream shear flow structure and altering the pressure distribution at the cavityʼs bottom.

Research Initiative for Numerical and Experimental Studies on High-Speed Stall of Civil Aircraft

The aerodynamics of transport aircraft at the borders of the flight envelope is characterized by complex interactions and nonlinear, unsteady flow phenomena. The underlying physical mechanisms are not fully understood, and a prediction of aerodynamic properties, transient loads, and aeroelastic behavior is a major challenge that requires the use of sophisticated numerical models and dedicated experiments. A new research initiative investigates the three-dimensional buffet mechanisms at high-speed stall, the interaction of the wing wake with the empennage, and the influence of ultra-high-bypass-ratio-type engines on these phenomena. It comprises unsteady pressure-sensitive paint surface pressure and time-resolved particle image velocimetry flowfield measurements carried out in the cryogenic European Transonic Windtunnel. The experiments are conducted at different Mach numbers and cover a broad Reynolds number regime up to using the generic Airbus XRF1 configuration as a reference. This paper explains the rationale of the research initiative, introduces the planned test entries and test conditions, and presents selected results of the first measurement campaign and associated numerical analyses by means of the computational fluid dynamics code TAU, including usage of locally scale resolving simulations. An analysis of numerical and experimental results is presented for a high incidence transonic condition that serves to discuss the topology of wing root flow separation, the evolution of the associated wake, and its interaction with the empennage.

Prediction of transonic wing buffet pressure based on deep learning

In the present study, a deep learning approach based on a long short-term memory (LSTM) neural network is applied for the prediction of transonic wing buffet pressure. In particular, fluctuations in surface pressure over a certain time period as measured by a piezoresistive pressure sensor, are considered. As a test case, the generic XRF-1 aircraft configuration developed by Airbus is used. The XRF-1 configuration has been investigated at different transonic buffet conditions in the European Transonic Wind tunnel (ETW). During the ETW test campaign, sensor data has been obtained at different local span—and chordwise positions on the lower and upper surface of the wing and the horizontal tail plane. For the training of the neural network, a buffet flow condition with a fixed angle of attack alpha and a fixed sensor position on the upper wing surface is considered. Subsequent, the trained network is applied towards different angles of attack and sensor positions considering the flow condition applied for training the network. As a final step, the trained LSTM neural network is used for the prediction of pressure data at a flow condition different from the flow condition considered for training. By comparing the results of the wind tunnel experiment with the results obtained by the neural network, a good agreement is indicated.

Aerodynamic Shape Optimization of a Symmetric Airfoil from Subsonic to Hypersonic Flight Regimes

Hypersonic flight has been the subject of numerous research studies during the last eight decades. This work aims to optimize the aerodynamic performance of a two-dimensional baseline airfoil (NACA0012) at distinct flight regimes from subsonic to hypersonic speeds. A mission profile has been defined, where four points representing the subsonic, transonic, supersonic, and hypersonic flow conditions have been selected. A framework has been implemented based on high-fidelity RANS computational fluid dynamics simulations. Gradient-based optimizations have been conducted with the objective of minimizing the drag. The optimization results show an overall improvement in aerodynamic performance, including a decrease in the drag coefficient of up to 79.2% when compared to the baseline airfoil. In the end, a morphing strategy has been laid out based on the optimal shapes produced by the optimization.

Investigation of Surge in a Transonic Centrifugal Compressor With Vaned Diffuser: Part II—Correlation With Subcomponent Characteristics

AbstractIn the first part of the paper, surge signatures and the underlying mechanisms thereof for a high-speed centrifugal compressor were investigated using signals from fast-response transducers installed along the flow path. The compressor surge signature is observed to vary as the impeller inlet tip flow transitions from subsonic to supersonic conditions. Spike-type deep surge is observed at subsonic and supersonic impeller inlet tip conditions while modal-type mild surge occurs at 90% speed with transonic inlet tip conditions. In Part II of the paper, a detailed analysis of the static pressure rise characteristics at the stage, component, and subcomponent levels is conducted. Additionally, the influence of the impeller inlet conditions on the instability is also considered. A connection between the surge mechanism and component static pressure rise characteristics is shown which exhibits the potential for prediction of the stall inception mechanism including identification of the destabilizing component. Furthermore, the transition from subsonic to transonic impeller inlet conditions is shown to be the cause of the unique mild surge instability at 90% speed. Lastly, both experimental and computational results are utilized to investigate the development of a shock wave at the impeller leading edge. The shock is considered to be critical to stage stability at speeds where the compressor experiences transonic inlet conditions.

The Role of Turbine Operating Conditions on Combustor–Turbine Interaction—Part I: Change in Expansion Ratio

Abstract Aeroengine lean-burn combustors release vorticity and temperature perturbations that, interacting with the first turbine stage, impact the stage aerodynamics, the blade cooling, and noise production. The first of these issues is addressed in this paper that is Part I of a two-fold contribution. A detailed experimental analysis is carried out to study the impact on the combustor–turbine interaction of the off-design conditions experienced by aero-engines in their duty. Engine-representative disturbances are generated by a combustor simulator able to produce swirling entropy waves. Two injection positions and four injection cases are studied. Experimental measurements are carried out at three traverses: upstream of the stator, at the interstage, and downstream of the rotor. This paper analyses the effect of the stage expansion ratio: two values are studied, namely 1.4 and 1.76, representative of subsonic and transonic flow conditions. They are chosen imposing similar velocity triangles at the rotor inlet. Results show that the swirl profile considerably impacts the stage aerodynamics. The aerothermal flow field downstream of the stator is modified significantly by the combustor disturbances. Conversely, downstream of the rotor, the differences in aerodynamics lessen. However, the entropy wave persists at the stage outlet and its transport depends on both the operating point and the injection position.

Investigation of Surge in a Transonic Centrifugal Compressor With Vaned Diffuser—Part I: Surge Signature

Abstract The identification of stall inception mechanisms and stability-limiting components in a centrifugal compressor is required for the development of effective surge suppression approaches. Part I of this two-part paper investigates the surge signature of a centrifugal compressor at subsonic, transonic, and supersonic inlet tip conditions, and Part II considers the relationship of the surge signature with the compressor static pressure characteristics, as well as the impeller leading edge relative tip Mach number. Experiments were performed in the single stage centrifugal compressor (SSCC) facility at Purdue University with both steady performance and dynamic pressure data recorded. Results show the presence of both mild and deep surges on the compressor map. Deep surge occurs at subsonic and supersonic impeller tip relative Mach numbers while the mild surge is observed as the tip relative Mach number nears unity. Long length-scale modal oscillations can be detected only at the transonic operating condition while spike disturbances are identified as the precursor to instability at subsonic and supersonic tips relative Mach numbers. Finally, the impeller is determined to be the origin of instability throughout most of the compressor map. The observation of impeller-induced instability refreshes the conventional understanding that the diffuser is typically the stability-limiting component for centrifugal compressors with vaned diffusers and suggests various surge suppression approaches may be necessary to achieve range extension throughout a compressor map.

Cooperation of Trailing-Edge Flap and Shock Control Bump for Robust Buffet Control and Drag Reduction

At transonic flight conditions, the buffet caused by the interaction between the shock waves and the boundary layers can degrade an aircraft’s aerodynamic performance and even threaten its safety. In this paper, the shock control bumps, originally designed to reduce the wave drag at cruise speeds, are applied to enhance the robustness of the closed-loop buffet control system using the trailing-edge flap. For the OAT15A supercritical airfoil, a closed-loop buffet control system is first designed with a feedback signal of lift coefficient. Then, the shock control bumps designed for drag reduction are integrated into the active buffet control system. The results show that the closed-loop flap control can be greatly enhanced by coupling with the shock control bumps. At the steady state under control, the shock control bumps can slightly increase the airfoil lift–drag ratio. More importantly, the ranges of control parameters that can effectively suppress the buffet are significantly enlarged with the help of the bumps; thus, the robustness of the control system is greatly enhanced.

On the Effect of Frequency Separation, Mass Ratio, Solidity, and Aerodynamic Resonances in Coupled Mode Flutter of a Linear Compressor Cascade

Abstract At a low mass ratio of structure to air, the work-per-cycle approach, or better known as the energy method, will lead to nonconservative results as aerodynamic coupling of modeshapes acts destabilizing. Using the p–k method to solve the aeroelastic stability equation, the effects of various structural aspects are investigated for a two-dimensional compressor cascade in subsonic and transonic flow conditions. The investigated key parameters are frequency separation, mass ratio, and solidity. Furthermore, the effect of a high frequency dependency of the aerodynamic forces is presented. Such phenomena can happen in case of aerodynamic or acoustic resonances. If the resonance peaks are close to the aeroelastic frequency, a discontinuous behavior of the frequency or damping solution can lead to a rapid destabilization of the system, once the aeroelastic frequency moves from one side to the other of the peak. In these regimes, it is crucial to have a high quality representation of the frequency-dependent generalized aerodynamic forces (GAFs) for an accurate prediction of the flutter onset.

Upstream and Passage Endwall Film Cooling of a Transonic Turbine Vane

Adiabatic film cooling effectiveness was measured on the endwall of a turbine vane under transonic flow conditions using binary pressure-sensitive paint. The combined effect of upstream and passage film cooling holes was evaluated. The experiments used a five-vane annular cascade in a blowdown wind tunnel. The mainstream velocity was set to exit isentropic Mach numbers of 0.7 and 0.9. The total coolant-to-mainstream mass flow ratio (MFR) was varied from 0.75 to 2.50%. Coolant-to-mainstream density ratios (DRs) of 1.0 and 2.0 were investigated. The first row was located upstream of the vanes, and the second and third rows were located in the passage between the vanes. Coolant was supplied by a divided plenum, allowing independent control of the MFR to the upstream and the passage holes. The coolant MFR was evenly split between the upstream and passage holes. Generally, increasing the MFR, the DR, and the Mach number increased the film cooling effectiveness. However, increasing the DR at MFRs of 0.625% and below decreased the laterally averaged film cooling effectiveness by as much as 60%, largely because mainstream flow was ingested through the second row. Upstream coolant provided some protection through the passage, with effectiveness as high as 0.2 at the second row.

Comparing Gradient-Free and Gradient-Based Multi-Objective Optimization Methodologies on the VKI-LS89 Turbine Vane Test Case

Abstract The present paper addresses the multi-objective aerodynamic shape optimization of the two-dimensional LS-89 turbine cascade. The objective is to minimize the entropy generation at subsonic and transonic flow conditions while maintaining the same flow turning. Nineteen design variables are used to parametrize the geometry. The optimization problem is used to compare two major classes of optimization algorithms and at the same time deduce if this problem has multiple local solutions or one global optimum. A first optimization strategy uses a gradient-based Sequential Quadratic Programming (SQP) algorithm. This SQP algorithm allows to directly handle the non-linear constraints during the optimization process. An adjoint solver is used for computing the sensitivities of the flow quantities with respect to the design variables, such that the additional gradient computational cost is nearly independent of the number of design variables. In addition, the same optimization problem is performed with a gradient-free-metamodel assisted-evolutionary algorithm. A comparison of the two Pareto-fronts obtained with both methods shows that the gradient-based approach allows to find the same optimum at a reduced computational cost. Moreover, the results suggest that the considered optimization problem is uni-modal. In other terms, it is characterized by a single optimal solution.

Dusting Hole Film Cooling Heat Transfer on a Transonic Turbine Blade Tip

Investigated is a transonic turbine blade tip with a squealer rim and a squealer recess, with a single dusting film cooling hole contained within the leading edge region of the squealer recess. Data are provided for transonic flow conditions for a range of film cooling blowing ratios for two tip gap values, using a linear cascade, with no relative motion between the blade and the casing. Surface heat transfer characteristics are measured using the transient impulse-response measurement approach, employed with infrared thermography. Line-averaged adiabatic film cooling effectiveness values, for the 1.4 mm tip gap, are generally very small along the pressure side rim, with only small, locally increased values along the suction side rim. For the 0.8 mm tip gap, line-averaged adiabatic film cooling effectiveness values are generally somewhat higher along the pressure side rim and along the suction side rim. In general, effectiveness values for both tip gap values, for these locations, and for the recess region, increase as the blowing ratio increases. As the tip gap decreases from 1.4 mm to 0.8 mm, line-averaged adiabatic film cooling effectiveness generally increases on the rims and downstream regions of the recess, with increased magnitudes which are spread over larger spatial surface areas. For tip gaps of 0.8 mm and 1.4 mm, for regions where the line-averaged heat transfer coefficient ratio deviates significantly from 1.00, values generally decrease as the blowing ratio increases. Across every region of the blade, line-averaged heat transfer coefficient ratios either decrease or remain approximately invariant, as the tip gap value decreases from 1.4 mm to 0.8 mm.

Effects of pressure side film cooling hole placement and condition on adiabatic film cooling effectiveness characteristics of a transonic turbine blade tip

The effects of film cooling hole placement location along the upper pressure side of a transonic squealer are considered. The thermal performance of four different film cooling configurations; B1, B2, B3 and B4, are considered using the University of Alabama in Huntsville's SS/TS/WT (supersonic/transonic/wind tunnel) experimental facility and a simulated turbine blade row using a linear cascade. Surface-varying results are provided for both the squealer blade tip surface, and for the upper pressure side of the squealer blade. These results are given for blowing ratios ranging from 0.42 to 3.20 in the form of spatially-resolved and spatially-averaged adiabatic film cooling effectiveness distributions. Because local static pressure variations and gradients vary in a significant manner at different blade locations, coolant accumulations near surfaces (and the associated surface thermal protection and film effectiveness distributions) produced by the different film cooling configurations are vastly different along the squealer blade tip surface. The B1 and B3 film cooling configurations show that substantial magnitudes of adiabatic film cooling effectiveness are present along the pressure side rim and the suction side rim. Also present is a considerable region of locally increased film cooling effectiveness and a wider region of surface protection within the squealer recess region for the B2 film cooling configuration. The contributions are unique because of a deficit of experimental data for film cooled squealer blades which operate with transonic flow, and because the present investigation is the first to consider the effects of hole placement location for upper pressure side film cooling arrangements with transonic flow conditions.

Evaluating the Aerodynamic Damping At Shock Wave Boundary Layer Interacting Flow Conditions with Harmonic Balance

Abstract The aerodynamic damping of a stator vane located in a rear stage of a high pressure compressor is evaluated at transonic flow conditions by numerical means. The results of a solution method based on temporal linearization around a steady RANS state are compared to results generated by relying on a state-of-the-art harmonic balance solver. It is found that the time-linearized method is not capable to reproduce the damping behavior in a sufficient manner for the majority of assessed nodal diameters. The limitations of the time-linearized method consisting in expansion around an imperfect steady RANS state, neglecting non-linear contributions and considering turbulence to be frozen at its steady state are evaluated and quantified. Furthermore, the presence of unsteady content induced by shock wave boundary layer interaction can be identified by performing a full-annulus URANS simulation based on time-integration. For a limited range of nodal diameters, this shock boundary layer interaction locks in to the assessed flutter motion and affects the damping behavior substantially. It is demonstrated that the mechanism of the shock wave boundary layer interaction can be reproduced with harmonic balance in accordance to the time-integration method. Coupling the eigenmode of interest and the identified shock boundary layer interaction via the harmonic balance method allows to predict the aerodynamic damping for the affected nodal diameters.

Analysis of tip leakage flow unsteadiness in a transonic turbine cascade using data-driven modal decomposition methods

The evolution of tip leakage flow and subsequent vortical structures is accompanied by inherent unsteadiness. This paper presents a novel characterization for the unsteady dynamics of turbine tip flow using data-driven, i.e., equation-free modal decomposition methods, which are applied to the hybrid Reynolds-averaged Navier–Stokes/large eddy simulation data at a transonic condition. By combining these techniques, the identified eigen-tuples (eigenvalues, eigenvectors, and time evolution) are well determined, and the differences between the obtained patterns (modes) are also pronounced. The snapshot proper orthogonal decomposition (POD) analysis can isolate the large-scale fluctuating structures that populate the rear part of the suction-side, which is mainly attributed to the shock-induced vortex instability. Similar to the turbulence cascade, macro-scale coherent structures that correspond to the tip leakage vortex shedding phenomena and the subsequently transitional and noisy parts closely related to the dissipation can be well derived by a quadruple reconstruction. Three dynamic mode decomposition (DMD) variants including the amplitude selecting-DMD method, the DMD with criterion method, and the sparsity promoting (SP)-DMD method are also compared in extracting dominant modes from the periodic tip flow, and the SP-DMD method which can distill modes of broadband frequencies and low dissipation is proved to be more conducive to representing and reconstructing the complex tip flow. Additional spectral-kernel-based POD (SPOD) analysis that can identify the similar primary unsteadiness frequencies as the DMD method is also encompassed in this study. Specifically, although it manifests that a physical resemblance of the pattern of pressure fluctuations to tip eddy unsteadiness can be captured by all these approaches, the behavior of small-scale vortical interaction downstream of the trailing edge can be clearly isolated with the intrinsic Karman-type vortex layer shedding process via DMD and SPOD approaches, which also demonstrates that these two techniques are more favorable to decomposing the complex tip flows into uncoupled single-frequency coherent structures compared to the conventional POD method. On this basis, resulting modes of velocity components have been accounted for verifying their contributions to the turbulent kinetic energy fields. The ensuing observations can offer a glimpse of the complex dynamics in the tip region, which also sheds light on features previously masked by conventional analysis approaches.

Martian Combustion-Powered Fixed-Wing UAVs: An Introductory Investigation and Analysis

The Martian topography needs to be investigated in greater detail for human habitations, and this can be accomplished faster using unmanned aerial vehicles (UAVs). In this regard, the RQ-11B Raven appears suitable for remote sensing and topography-mapping applications on Mars, due to its popularity in surveillance and reconnaissance applications on Earth. As a result, this study investigates the flight of this UAV in the Martian atmosphere with the assumptions that it employs an NACA S7012 airfoil and its electric propulsion technology is replaced with a four-stroke oxy-methane fueled Saito FG-11 internal combustion engine (ICE). This ICE is estimated to supply 367.8 W resulting in an engine speed of 6891 revolutions per minute. Based on this speed, the UAV must fly at least 72 m/s (Re = 18,100) at a 5° angle of attack to support flight under calm conditions. To achieve this speed will be difficult; thus, a weather balloon or German V1-style launch system should be employed to launch the UAV successfully. Furthermore, the UAV must operate below 165 m/s (Re = 41,450) to prevent transonic conditions. Finally, the vehicle’s fuel and oxidizer tanks can be refueled using an in situ methane and oxygen production system, enabling its sustainable use on Mars.

Turbulent flow simulations of the common research model on Cartesian grids using recursive fitting approach

In this study, we propose a new method that uses the recursive fitting method and wall function for the three-dimensional turbulent flow simulation. The recursive fitting method based on the Cartesian grid method is employed for the automatic generation around two- and three-dimensional geometries. The grid approximately represents the concave and convex features, i.e., a wing–fuselage juncture and wing trailing edges. The wall boundary conditions can be imposed on the wall faces; thus, the conservation laws are strictly satisfied. Meanwhile, using the wall function allows turbulent boundary layers to be accurately reproduced. The turbulent flow simulation using the wall function is straightforward because the flow variables at the first cell centroid are employed for friction velocity estimation. However, the distance between the body surface and the first cell centroid varies since cells with different shapes are generated. Therefore, the simulation using the proposed method is validated through turbulent flow simulations in two dimensions. The results show the accurate predictions of the turbulent boundary layers. To investigate the capacity of the proposed method, three-dimensional simulation around aircraft at a transonic cruise condition is conducted. The automatically generated grids by the proposed method approximately represent the geometric features. The results of flow simulation show that the surface pressure and skin friction coefficient distributions agree with those of the reference simulation on a conventional body-fitted grid. Further, predicted pressure and viscous drag coefficients are reasonable compared with those of the reference simulation.

Uncertainty Quantification of Multivariate Correlated Output in the Simulation of Airfoil Flow

Due to the limitation of human cognition, there are uncertainties in the turbulence model closure coefficients that cannot be ignored. The key to assess model objectively is quantifying the uncertainty of model coefficients scientifically. In the CFD simulation of flow around the airfoil, wall pressure distribution is a multivariate correlated quantity. This paper developed a surrogate model method based on artificial neural network to characterize the mapping relationship between multivariate input and multivariate output. In order to reduce the output dimension and modelling difficulty, this paper use Proper Orthogonal Decomposition (POD) method. An RAE2822 benchmark case under transonic conditions is used to study the uncertainty of Spalart-Allmaras (SA) turbulence model coefficients and verify the method developed by this paper.

Global aero-structural design optimization of composite wings with active manoeuvre load alleviation

In the scope of the DLR project VicToria (Virtual Aircraft Technology Integration Platform), an integrated process for aero-structural wing optimization based on high fidelity simulation methods is continuously developed and applied. Based upon a parametric geometry, flight performance under transonic flight conditions and manoeuvre loads are computed by solving the Reynolds-averaged Navier–Stokes equations. Structural mass and elastic characteristics of the wing are determined from structural sizing of the composite wing box for essential manoeuvre load cases using computational structural mechanics. Static aeroelastic effects are considered in all flight conditions and active manoeuvre load alleviation is integrated in the process. Global aero-structural wing optimizations are successfully performed for wings with and without active manoeuvre load alleviation. The active manoeuvre load alleviation is introduced with a simplified modelling of control surface deflections using a mesh deformation technique. The minimization of the fuel consumption for three typical flight missions represents the objective function. Wing optimizations are performed for variable and constant wing planform parameters as well as for wings with conventional composite wing box structure and for more flexible wings. The latter is accomplished by introducing modifications of the structural concept and the strain allowable. A significant mass reduction of the optimized wing box is obtained for wings with active manoeuvre load alleviation, resulting in a drop in fuel consumption of about 3%. For wing optimizations with the more flexible wing concept, the active manoeuvre load alleviation shows an additional reduction of the fuel consumption in the order of 2%. The wings with active manoeuvre load alleviation results in optimized wing geometries with increased aspect ratio and reduced taper ratio.