Subsonic Conditions Research Articles

Influence of strut on trapped vortex cavity flameholder: Ignition and flame propagation performances

In order to improve the performance of pilot flame propagation to the core flow, a novel strut design was proposed for the trapped vortex cavity flameholder. Experiments were conducted under high subsonic flow conditions to investigate the influences of strut on the ignition performance and flame propagation performance of trapped vortex cavity combustor. Furthermore, the flame development was acquired using a high-speed camera and the flame information and the flow field were combined to explicate the flame stabilization mechanisms of two flameholder layouts. The results showed that, both flameholder layouts achieved reliable ignition and flame stability under inlet flow velocity lower than Ma = 0.42. Cavity combined with strut weakened the ignition performance of the trapped vortex cavity-only flameholder based combustor, but improved the blowout performance and enhanced the flame stability. Compared with the cavity-only flameholder layout, the lean ignition equivalence ratio of the cavity-strut flameholder layout was increased of 14.6%, and the lean blowout equivalence ratio was decreased by 56.2%, under the inlet flow at 600 K. Strut changed the original flame propagation path of the cavity-only layout, and caused the flame stabilization mechanism change from cavity flame stabilization to strut flame stabilization. Although the change of flame stabilization mechanism prolongs the ignition delay time, it enhances the flame stability of the combustor. Moreover, the flame propagation ability to the radial direction behind the cavity combined with strut is enhanced, which is more suitable to the engineering combustor with higher requirements on circumferential flame propagation.

Modelling aileron and spoiler deflections with the linear frequency domain method (LFD) for subsonic flight conditions

PurposeThis study aims to predict dynamic responses of aileron and spoiler control surfaces in subsonic flight via the use of surrogate models. The prepared reduced order models prove useful when quick estimations for a large number of variations are required.Design/methodology/approachThe linear frequency domain (LFD) method was used for the simulation study. Each surrogate contained a database of 100 control surface dynamic responses over a spectrum of 200 harmonics computed with LFD. To interpolate new results, the DLR surrogate modelling toolbox, SMARTy, was used. The database’s samples were prepared in a Halton sequence, making interpolation reliable. The surrogate’s parameter space was the Mach number, Reynold’s number, angle of attack, control surface deflection angle and the control surface chord length.FindingsThe LFD method proved effective for the mentioned purpose: the surrogates were accurate, up to 15% of relative error, in reproducing dynamic responses of aileron and spoiler deflections at low speed, within the limitations of flow field linearity, as well as surrogate prediction capability. The restrictions of the surrogate, and the reasoning thereof, are also presented in detail in the study. Future load alleviation studies are a potential of the findings here.Originality/valueLFD is an innovative technique for load prediction and alleviation studies. This paper provides a reference for engineers wishing to use the method for the two mentioned control surfaces, or the like.

CFD analysis of flow control in compressor cascade using MVGs

Abstract The present study reports the numerical investigation of the compressor cascade. To minimize the separation phenomenon in the compressor cascade, a passive flow control device i.e. Micro Vortex Generator (MVG) is utilized. MVG is a very simple and lightweight attachment mounted infront of the leading edge of the cascade blade. Due to being passive in nature, it neither consumes power nor requires any external device to actuate. The numerical simulations were carried out on a highly loaded compressor cascade at an angle of incidence of −1° under subsonic conditions at Mach number 0.2. The profile of the compressor cascade blade was double circular arc (DCA), unsymmetrical and cambered at 40°. Two different types of micro vortex generators were mounted infront of the leading edge in the compressor cascade to control the secondary flows since secondary flows were responsible for various losses in cascade. To analyze the flow under incompressoible state of air (M = 0.2), Star CCM + software has been used. To simulate the flow under turbulent condition, k-ω SST turbulence model was used. A velocity profile of 25 mm boundary layer thickness was extracted and used as an input in the compressor cascade. Mounting of MVG on compressor cascade enhanced drag but also increased lift. Total pressure loss coefficient (TPLC) was calculated to compare the losses. The aerodynamic efficiency in terms of coefficient of lift and coefficient of drag has been used to study the effect of MVG over cascade. It is found that there is reduction in total pressure loss coefficient (TPLC) for trapezoidal and curved trapezoidal types of MVGs and the decrease in percentage are 2.17 and 8.86%, respectively. Alos, aerodynamic efficiency is increased by mounting trapezoidal and curved trapezoidal types of MVG and the increase in percentages are 2.03 and 3.10%, respectively.

On the Rossiter-Heller frequency of resonant cavities

The goal of this paper is to revise the Rossiter-Heller formula of cavity resonance frequencies for rectangular cavities in subsonic and supersonic flow conditions. Two modifications are proposed; the first one consists in using a second-order polynomial fit of literature experimental data for the non-dimensional time lag coefficient as a function of the cavity aspect ratio. The second modification consists in taking into account the reversed flow convective effects in the calculation of the backward acoustic propagation time, by including the reversed flow Mach number in the formula, expressed as a function of the free-stream Mach number and cavity aspect ratio. A second-order polynomial fit of scale-resolving flow simulation results of the reversed flow Mach number is used to derive the final resonance frequency formula. The effects of the two modifications are investigated for three values of the aspect ratio in the range 5 to 10, and four values of the free-stream Mach number in the range 0.6 to 1.35. It is concluded that taking into account both modifications yields globally more accurate predictions of the resonance frequencies. Finally, an attempt is made to identify the nature of additional tones observed for the deep supersonic case around the fourth and fifth Rossiter-Heller resonance modes.

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.

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.

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 II: Loading Effects

Abstract Aero-engine combustors burn a lean and premixed blend releasing vorticity and temperature perturbations. Interacting with the first turbine stage, these disturbances impact the cascade aerodynamics, add criticality to the blade cooling, and are sources of noise. The first of these issues is addressed in this paper, focusing on off-design turbine conditions, as experienced by aero-engines in their duty. This paper, Part II of a two-fold contribution, analyses the effect of the stage loading obtained by changing the rpm (three different values) at the same expansion ratio of 1.4, representative of subsonic flow conditions. Engine-representative disturbances are generated by a combustor simulator able to produce a swirling entropy wave. Two injection positions and four injection patterns are considered. Experimental measurements are carried out through the stage, measuring the injected disturbance and the aerothermal flow field downstream of the stator and the rotor. Results show that the swirl profile mostly impacts the stage aerodynamics. The different work extraction and the interaction with secondary flow structures change the entropy wave transport, diffusion, and decay through the rotor. Furthermore, the increased angle of the incidence caused by the injected disturbance can make the blade stall under the most loaded operating condition.

A wide-speed-range aerodynamic configuration by adopting wave-riding-strake wing

The wide-speed-range characteristics of aerodynamic configuration are critical to the design of the horizontal take-off and landing aircraft. To address this problem, the work proposes a wide-speed-range aerodynamic configuration by assembling a waverider to the aircraft with arrow-shaped wing. To enhance the aerodynamic performance of this configuration, surrogate optimization method is applied to the airfoil of the arrow-shaped wing. Thus, the optimized airfoil with double “S” cambers on the lower surface, and a small leading-edge radius is obtained. The aerodynamic performance of the whole aircraft in the wide-speed range has been investigated numerically by means of the three-dimensional Reynolds-averaged Navier-Stokes (RANS) equations coupled with the SST k-omega turbulence model. At subsonic conditions, the waverider can act as a strake-wing to induce vortex from the leading edge of the upper surface. With the large sweep arrow-wing, the subsonic aerodynamic performance is improved by utilizing the vortex lift. At supersonic conditions, wave-riding characteristics as well as the small leading-edge radius wing can effectively reduce the wave-drag. Overall, the lift-to-drag ratio of this configuration is not less than 5.4 at 4-degree angle of attack within Mach number 0.5–7, indicating that this configuration provides outstanding hypersonic aerodynamic performance as well as fairly acceptable subsonic aerodynamic characteristics.

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.

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.

Laminar flame speed and autoignition characteristics of surrogate jet fuel blended with hydrogen

Hydrogen combustion has emerged as one promising option toward the achievement of carbon-neutral in aviation. In this study, the effects of hydrogen addition on laminar flame speeds, autoignition, and the coupling of autoignition and flame propagation for surrogate jet fuel n-dodecane are numerically investigated at representative engine conditions to elucidate the potential challenges for flame stabilization and the autoignition risks in combustor design. Results show that the normalized flame speed increases almost linearly with hydrogen addition for fuel-lean conditions, while for fuel-rich conditions it increases nonlinearly and can be up to 20. This poses great challenges for avoiding flameholding and flashback, particularly for fuel-rich mixtures. Results further show that flame speed enhancement due to the increased flame temperature can be neglected under fuel-lean conditions, but not for fuel-rich mixtures. For the dependence of ignition delay time on temperature, there exists a unique intersection between pure n-dodecane/air and H2/air mixtures. Near the intersection temperature, there exists subtle kinetic coupling of the two fuels, leading to different H2 roles, e.g., accelerator or inhibitor, for the autoignition process of n-dodecane/H2/air mixtures. With this intersection temperature, the diagram for autoignition risks is constructed, which demonstrates that H2 acts as an inhibitor under subsonic cruise conditions while either an inhibitor or an accelerator under supersonic cruise conditions depending on the combustor inlet temperature and the amount of hydrogen addition. With the potential coupling of autoignition and flame propagation, the 1-D autoignition-assisted flame calculations show that hydrogen addition can alleviate or even eliminate the two-stage ignition characteristics for pure n-dodecane/air flames. For n-dodecane blended with hydrogen, the autoignition-assisted flame propagation speed, as well as the global transition from flame propagation to autoignition, can still be described by an analytic scaling parameterized by the ignition Damkӧhler number.

Effect of compound lean blades on separation structures in high loaded compressor cascades under high subsonic condition

Three-dimensional separation is an inherent flow feature in the blade-end corner of compressors passage, which is a primary source of entropy generation and loss. This paper researches the effect of key geometrical parameters (camber angle, solidity, dihedral angle of compound lean blade) and incoming condition (Mach number) on the evolution of corner separation in a low aspect ratio linear cascade. The underlying flow mechanism is explored in detail. The evolution of typical flow characteristics with the variation of inlet Mach number, camber angle and solidity is interpreted. For different blade loading levels, there are different corner separation forms and vortex structures. The influence of dihedral angle on the flow field structure and flow loss is analyzed, thus the effect mechanism of the compound lean blade in different flow environment is explored. Without the trailing edge shedding vortex, positive dihedral angle alleviates the low momentum fluid accumulation and corner separation at the incidence angle about 0°, but exacerbates the deteriorating corner separation at large positive incidence angle. When the trailing edge shedding vortex and the suction surface separating vortex exist together, positive dihedral angle promotes the upstream migration of the trailing edge shedding vortex, which helps to truncate the suction surface separating vortex. This variation of vortex structures is conducive to weaken the development of corner separation and delay corner stall.

Preliminary Design and Analysis of Supersonic Business Jet Engines

Currently projected supersonic business jets target selected supersonic flight missions with Mach numbers of about 1.4 and a larger number of long-range subsonic flight missions. They form a new type of aircraft that is specially tailored to these requirements. The question arises as to which engine configurations and technology levels are required to support these new applications. This is addressed firstly by exploring the design space of potential working cycles. An aircraft model is used to translate the results of the cycle study into an expected aircraft range. An optimal core engine and fan configuration result from the cycle study and the derived mission ranges. The preliminary design of the low-pressure components is investigated in the second step based on the optimal core configuration. The highest non-dimensional parameters are encountered in subsonic flight conditions. The highest dimensional parameters are encountered in supersonic high-altitude flight conditions. High-overall-efficiency configurations do not result in optimal aircraft ranges. There is an optimal number of two fan stages and a specific thrust of about 300 m/s, resulting in a maximum aircraft range that is 11% superior to that achievable with a single-stage fan. A fan hub-to-tip ratio range that is comparable to that of military fans is desirable, with an aerodynamic lower limit around 0.37. The low-pressure turbine stage count is a compromise between turbine mass and size.

Progress and challenges in exploration of powder fueled ramjets

• A detailed review of the research progress on powder fueled ramjets is presented. • The key technologies and scientific issues in powder fueled ramjets are proposed. • Some suggestions are proposed for the development of powder fueled ramjets. The powder fueled ramjet is a developing propulsion system, which makes use of metal powder as the propellant and presents noticeable advantages, for instance, high specific impulse and density-specific impulse, thrust adjustment, simple structure, and suitability for high Mach flight vehicles. In this paper, the basic work characteristics are introduced and the exploration progress is reviewed and summarized for powder fueled ramjets. Furthermore, several key technologies are proposed, which mainly include the powder feeding system, ignition and combustion processes of the powder fuel, combustion organization technology, multiple start, thrust adjustment, and thermal protection technology. The development of the powder fueled ramjet is in the proof-of-concept stage, and a large number of individual technology breakthroughs have been achieved in subsonic conditions. However, in both subsonic and supersonic conditions, the combustion mechanism of powder fuel needs to be investigated by high precision gas–solid two-phase numerical simulation methods, and ground tests, thus, to guide combustion chamber design. In order to achieve thrust regulation, coupling studies on powder fuel delivery and combustion are required. Besides, new thermal protection schemes based on the lattice structural and supercritical fluid active cooling have great potential to meet the demands of complicated two-phase combustion thermal environments.

High subsonic and low reynolds number airfoil optimization of high-altitude propeller

This paper presents an airfoil optimization at high subsonic and low Reynolds number condition, to improve the efficiency of an E387-based HALE UAV propeller at cruising condition. The optimization design method is based on surrogate model coupled with validated RANS simulation. The optimal airfoil shape is flat and sharp leading-edged, which remarkably differs from the original and resembles a very thin airfoil in reversed configuration. Significant increment of lift-to-drag ratio and aerodynamical independency of Mach and Reynolds number were observed due to expanded supersonic flow region, weakened shock wave, and considerably reduced flow separation. By applying the optimized airfoil to propeller outboard, the new propeller’s efficiency increased more than 2%, while maintained required thrust, as proved by both Blade Element Theory and CFD simulation.

Numerical Investigation of the Aerofoil Aerodynamics with Surface Heating for Anti-Icing

The aerodynamics of an aerofoil with surface heating was numerically studied with the objective to build an effective anti-icing strategy and balance the aerodynamics performance and energy consumption. NACA0012, RAE2822 and ONERA M6 aerofoils were adopted as the test cases and the simulations were performed in the subsonic flight condition of commercial passenger aircraft. In the first session, the numerical scheme was firstly validated with the experimental data. A parametric study with different heating temperatures and heating areas was carried out. The lift and drag coefficients both drop with surface heating, especially at a larger angle of attack. It was found that the separation point on the upper surface of the aerofoil is sensitive to heating. Higher heating temperature or larger heating area pushes the shock wave and hence flow separation point moving towards the leading edge, which reduces the low-pressure region of the upper surface and decreases the lift. In the second session, the conclusions obtained are applied to inform the design of the heating scheme for NACA0012. Further guidelines for different flight conditions were proposed to shed light on the optimisation of the heating strategy.

A Grid Fusion Lifting Surface and Its Flow Control Mechanism at High Angles of Attack

A grid fusion lifting surface suitable for large subsonic angles of attack was designed. The influence of the grid position, grid diversion angle, and grid number on the aerodynamic performance of the grid fusion lifting surface under subsonic conditions was studied by numerical simulations. Based on the summary and analysis of the grid in the hypersonic state, the CFD numerical simulation method, which solves unsteady Reynolds averaged Navier-Stokes equations, was used to complete the research of the flow mechanism of the grid fin in the subsonic state. At low speed and subsonic conditions, the results show that the negative front grid fusion lift surface has better aerodynamic characteristics with a high angle of attack. Under subsonic conditions, the stall angle of attack of the front grid fusion lifting surface with a diversion angle of 20° increases by 16°, and the maximum lift coefficient increases by 22.1% compared to that of the conventional flat wing. The number of grids has an insignificant effect on the aerodynamic performance of the grid fusion lifting surface at a large angle of attack.

Fast Predictions of Aircraft Aerodynamics Using Deep-Learning Techniques

The numerical analysis of aerodynamic components based on the Reynolds–averaged Navier–Stokes equations has become critical for the design of transport aircraft but still entails large computational cost. Simulating a multitude of different flow conditions with high-fidelity methods as required for loads analysis or aerodynamic shape optimization is still prohibitive. Within the past few years, the application of machine learning methods has been proposed as a potential way to overcome these shortcomings. This is leading toward a new data-driven paradigm for the modeling of physical problems. The objective of this paper is the development of a deep-learning methodology for the prediction of aircraft surface pressure distributions and the rigorous comparison with existing state-of-the-art nonintrusive reduced-order models. Bayesian optimization techniques are employed to efficiently determine optimal hyperparameters for all deep neural networks. Three data-driven methods which are Gaussian processes, proper orthogonal decomposition combined with an interpolation technique, and deep learning are investigated. The results are compared for a two-dimensional airfoil case and the NASA Common Research Model transport aircraft as a relevant three-dimensional case. Results show that all methods are able to properly predict the surface pressure distribution at subsonic conditions. In transonic flow, when shock waves and separation lead to nonlinearities, deep-learning methods outperform the others by also capturing the shock wave location and strength accurately.