Pitching Moment Coefficient Research Articles

Analysis of blended winglet parameters on the aerodynamic characteristics of NXXX aircraft using Computational Fluid Dynamics (CFD)

Winglet is a wingtip device that was developed to improve aircraft flight performance by reducing induced drag. The induced drag produced by the NXXX aircraft is considered too excessive because the simple wingtip currently available is still not optimal in reducing induced drag, so a new more optimal wingtip is needed. This research aims to analyze the influence of blended winglet parameters on the aerodynamic characteristics of the NXXX aircraft such as CL and CD. Not only that, pitch moment coefficient (CM) is also studied to determine how trim drag can increase with the consequent increase of aerodynamic efficiency (CL/ CD) which currently has not been specifically researched yet. Winglet parameters in this study include variations in taper ratio, winglet height, cant angle, trailing edge sweep, and blending radius that there is not enough analysis about it now. Computational Fluid Dynamics (CFD) simulation uses Ansys CFX with the Shear Stress Transport (SST) turbulence model is used to obtain wing aerodynamic data. This research concludes that the best winglet configuration is taper ratio of 0.2, blending radius of 15 %, winglet height of 30 %, cant angle of 15°, and trailing edge sweep of −0.6°. Apart from that, it was also found that all winglet configurations increased CL, CLmax, and also lift slope (a). Almost all winglets also reduce the critical angle of attack, reduce CD, and increase CM. This simulation finds that addition of blended winglet can increase aerodynamic efficiency up to 17.51 % in cruise condition using the variation of winglet height.

Numerical Analysis on The Aerodynamic Performance of A High-speed Train Operating on Various Embankment Heights Under The Influence of Crosswinds

Given the unavoidable geographical surface, railings must be raised above ground level in some cases, which is known as an embankment. It was discovered that the height of the embankment had a significant influence on the slipstream on the train's leeward side especially during crosswind conditions. The primary objectives of this study were to investigate the impact of varying embankment heights on the aerodynamic characteristics of a high-speed train under different crosswind conditions using computational fluid dynamics (CFD) analysis. The German Aerospace Center DLR's Next-Generation High-speed Train (NG-HST) model has been used for this study. The yaw angles (ψ) are ranging from 10° to 50° in 10° increments. The Reynolds number based on the model's height and freestream velocity at the computational domain is 1.3 x 106. In the results, it shows that embankment height and crosswind ψ has a significant impact on the aerodynamic characteristics. Essential aerodynamic parameters that have a significant impact on train stability, such as the drag, lift, and side force coefficients, as well as the roll, yaw, and pitch moment coefficients, revealed that the higher the ψ, which included 40° and 50°, produced poor results compared to the lower ψ, which included 10°, 20°, and 30°. In terms of visual appearance, rising of crosswind angles have a greater impact on the formation of vortices on the leeward side of the train body and embankment. Thus, it can be concluded that the embankment heights and crosswind angles are crucial in determining train safety operations.

Nonlinear Surrogate Model Design for Aerodynamic Dataset Generation Based on Artificial Neural Networks

In this work we construct a surrogate model using artificial neural networks (ANN) to predict the steady-state behavior of an unmanned combat aircraft. We employ various strategies to improve the model’s accuracy, including the consideration of design tolerances, creating independent surrogate models for the different flow regimes and encoding non-numeric input features. We also explore alternative machine learning models, albeit they demonstrated a lower reliability than ANNs. Two scenarios are considered for the target variable: one focusing solely on predicting the pitching moment coefficient, and the other incorporating the roll moment coefficient as well. We investigate different methods for handling multiple targets, finding that constructing a single model with multiple outputs consistently outperforms developing separate models for each target variable. Overall, the ANN provides predictions that show excellent agreement with the experimental data, demonstrating its effectiveness and reliability in aerodynamic modeling.

Effects of the gap ratio on the flow field structures and the aerodynamic performance of an airfoil with ridge ice

Under SLD icing conditions, the ridge ice may appear on the surface of aircraft, which led to the significant aerodynamic deterioration and affected aircraft flight safety. The present study experimentally investigated the effects of the gap ratio on the flow field structures and aerodynamic performance of an airfoil with ridge ice. Detailed measurements were performed in a low-speed reflux wind tunnel by utilizing the Particle Image Velocimetry (PIV) technique and a high-sensitivity six-component balance. The results showed that the maximum lift coefficient, stall angle, and maximum pitch moment coefficient of the airfoil increased as the gap ratio enlarged. At AOA = 10 deg, the separation bubble length decreased by 77 % when the gap ratio changed from 0 to 0.1. Meanwhile, the separation bubble length decreased by 68 % when the gap ratio changed from 0.1 to 0.3. Besides, as the increase of the gap ratio, the average vorticity, turbulent kinetic energy, and Reynolds shear stress in the selected region above the airfoil decreased, while the average velocity increased. In addition, the gap ratio did not have an apparent effect on the transition onset positions and the maximum spanwise vorticity in the flow field.

Experimental Investigation on the Aerodynamic Parameters of Trailing Edge Wake Generators

In the realm of aerodynamics, the investigation of airfoil performance stands as a critical domain, with an ever-growing emphasis on optimizing designs for enhanced efficiency. This study investigates the aerodynamic performance of a NACA 0015 airfoil featuring various trailing edges, including serration, comb, poro-serration, and comb-serration. The experiments were conducted in a wind tunnel at angles of attack ranging from -15° to 15° and Reynolds numbers of 1.5 x 105 and 2.0 x 105. To accurately quantify the forces and moments acting on the airfoil models, a calibrated six-component balance was utilized to measure the aerodynamic coefficients of each airfoil configuration. The lift coefficient (cl), drag coefficient (cd), and pitching moment coefficient (cm) are analyzed for the various trailing edge types and Reynolds numbers. Results indicate that the baseline model demonstrated better aerodynamic performance compared to other types of trailing edge. Most trailing edges, except the baseline, resulted in a decrease in the lift coefficient. However, at very low angles of attack, the airfoil showed an improvement in the maximum lift coefficient. Most trailing edges exhibited an increase in the drag coefficient at a Reynolds number of 1.5 x 105. However, at a Reynolds number of 2.0 x 105, the drag coefficient showed a similar trend as the baseline. All types of trailing edges, including the baseline, displayed a similar trend in the pitching moment coefficient. When evaluating the lift coefficient to drag coefficient ratio, all trailing edges generally performed similarly at all Reynolds numbers. In general, the baseline model emerges as the optimal choice, showcasing superior aerodynamic characteristics across the evaluated parameters.

Fast fluid–structure interaction simulation method based on deep learning flow field modeling

The rapid acquisition of high-fidelity flow field information is of great significance for engineering applications such as multi-field coupling. Current research in flow field modeling predominantly focuses on low Reynolds numbers and periodic flows, exhibiting weak generalization capabilities and notable issues with temporal inferring error accumulation. Therefore, we establish a reduced order model (ROM) based on Convolutional Auto-Encoder (CAE) and Long Short-Term Memory (LSTM) neural network and propose an unsteady flow field modeling method for the airfoil with a high Reynolds number and strong nonlinear characteristics. The attention mechanism and weak physical constraints are integrated into the model architecture to improve the modeling accuracy. A broadband excitation training strategy is proposed to overcome the error accumulation problem of long-term inferring. With only a small amount of latent codes, the relative error of the flow field reconstructed by CAE can be less than 5‰. By training LSTM with broadband excitation signals, stable dynamic evolution can be achieved in the time dimension. CAE-LSTM can accurately predict the forced response and complex limit cycle behavior of the airfoil in a wide range of amplitude and frequency under subsonic/transonic conditions. The relative errors of predicted variables and integral force are less than 1%. The fluid–structure interaction framework is built by coupling the ROM and motion equations of the structure. CAE-LSTM predicts the time series response of pitch displacement and moment coefficient at different reduced frequencies, which is in good agreement with computational fluid dynamics, and the simulation time savings exceed one order of magnitude.

Six-dimensional force/torque sensor for aerodynamic characteristic study of high-speed train with different wind angles under stationary tornado.

Researchers conducted an investigation by tornado simulator to study the impact of wind angle on the aerodynamic characteristics of a reduced (1:150) high-speed train model using six-dimensional force/torque sensor. The reduced scale model size can match the relative size relationship between high-speed train and tornado vortex core in real condition. Results show that the wind angle affects the mean value and standard deviation of the force and moment coefficient of the high-speed train at the same radial position. The variations of the mean value and standard deviation of the pitching moment coefficient of the high-speed train carriage model at 60°and 90°are different from that at other wind angles. The variations of the mean value of the pitching moment coefficient of the high-speed train head model at 0°, 15°and 30°are different from that at other wind angles. The variations of the standard deviation of the pitching moment coefficient of the high-speed train head model at 60°,75°and 90°are different from that at other wind angles. This research will help the further study of the operation safety of high-speed train in the event of a tornado impacting a high-speed train network.

Liquid film morphology and aerodynamic performance of a high-speed train under wind-rain environments

The performance degradation of railway vehicles operating in wind-rain environments has attracted widespread attention in recent years. This work focuses on a numerical study of the liquid film morphology and train aerodynamic performance under wind-rain environments. The train aerodynamic model in wind-rain environments is set up using Eulerian–Lagrangian methodology, and validated with existing experimental results. The liquid film morphology of the train is discussed using the Lagrangian wall film (LWF) model. The aerodynamic characteristics such as liquid film morphology, surface pressure, shear stress, velocity contour, aerodynamic coefficient, etc., are then discussed in detail. The results show that, when the rainfall duration is constant, the liquid film becomes thicker as rainfall intensity is greater. As crosswind becomes stronger, the thickness of liquid film gathered on the train roof decreases, and the liquid film on the windward and leeward sides of the train becomes thicker. The shear stress on the train in wind-rain environments is obviously greater than that exposed to crosswind environments. For the reason of rainfall, the skin friction coefficient on the train also increases, which will have an impact on the frictional resistance. Rainfall would lead to an increase in the aerodynamic coefficient of trains operating in crosswind environments, especially for lift force coefficient and pitch moment coefficient.

Fixed-time angle of attack constrained control for aircraft considering dynamic icing process

Aircraft icing deteriorates aerodynamic performance and reduces stall angle of attack, the fast convergence rate of tracking error is required to stabilize the aircraft when aircraft icing occurs. The state-of-the-art control methods for icing aircraft mostly assume that the icing of aircraft is instantaneous. Aiming at these issues, a fixed-time angle of attack-constrained control strategy is designed considering dynamic icing process. In order to explore the variation of aerodynamic coefficients in the process of dynamic icing, an ice wind tunnel experiment is implemented, and the relationship between lift coefficient, drag coefficient and pitching moment coefficient with angle of attack and icing intensity is obtained by fitting method. In order to prevent the stalling problem caused by the decrease of the stalling angle of attack in the process of dynamic icing, a method to determine the stalling angle of attack based on deep neural network is proposed. Considering the asymmetric and time-varying angle of attack constraint, a fixed-time convergent angle of attack-constrained robust control method is designed. The ice wind tunnel experiment shows the process of dynamic icing of the airfoil, and the simulation results verify the effectiveness of the proposed control method.

Comparison of force measurement techniques in a short duration hypersonic facility

Force measurement experiments have been conducted within the University of Oxford’s High Density Tunnel with a 7∘\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$7^{\\circ }$$\\end{document} half-angle cone. The purpose of the study was to provide a direct comparison between two independent force techniques in the same facility at the same free-stream conditions to provide a quantitative and qualitative discussion of the advantages and disadvantages of both techniques. The first force measurement technique used a conventional 4-axis sting-mounted force balance which was calibrated both statically and through the stress wave deconvolution method, whilst the second technique used the less established static free-flight methodology. Experiments were conducted at a Mach 5 test condition which provided sufficient dynamic pressure to generate aerodynamic forces suitable for the measurement range of the force balance. Results for lift, drag and pitching moment coefficients were obtained over a range of angles of attack and compared with predictions from a hypersonic panel method code. Agreement between the independent force techniques and numerical data sets was good over the range of angles of attack. Maximum uncertainties were shown to be 38.46 ± 0.56 N and 22.52 ± 0.44 N for free-flight in lift and drag, respectively, and 38.74 ± 1.59 N and 22.13 ± 1.27 N for the dynamically calibrated force balance which demonstrates the superiority of free-flight.

Numerical Investigation on the Evolution Process of Different Vortex Structures and Distributed Blowing Control for Dynamic Stall Suppression of Rotor Airfoils

The influencing characteristic for the evolution mechanism of a dynamic stall vortex structure and distributed blowing control on rotor airfoils was investigated. Based on the moving-embedded grid method, the finite volume scheme, and Roe’s FDS scheme, a simulation method for the unsteady flow field of a pitch-oscillating airfoil was established. The flow field of the NACA63-218 airfoil was calculated using Reynolds-averaged Navier–Stokes equations. The evolution processes of different vortex structures during dynamic stall and the principal controlled vortex mechanism affecting aerodynamic nonlinearity were analyzed based on the pressure contours Cp and Q of the flow field structure and the spatiotemporal evolution characteristics of the wall pressure distribution. The research indicated that dynamic stall vortices (DSVs) and shear layer vortices (SLVs) were the major sources of the increase in aerodynamic coefficients and the onset of nonlinear hysteresis. Building upon these findings, the concept of distributed blowing control for DSVs and shear layer vortices (SLVs) was introduced. A comparative analysis was conducted to assess the control effectiveness of dynamic stall with different blowing locations and blowing coefficients. The results indicated that distributed blowing control effectively inhibited the formation of DSVs and reduced the intensity of SLVs. This led to a significant decrease in the peak values of the drag and pitch moment coefficients and the disappearance of secondary peaks in the aerodynamic coefficients. Furthermore, an optimal blowing coefficient existed. When the suction coefficient Cμ exceeded 0.03, the effectiveness of the blowing control no longer showed a significant improvement. Finally, with a specific focus on the crucial motion parameters in dynamic stall, the characteristics of dynamic stall controlled by air blowing were investigated. The results showed that distributed air blowing control significantly reduced the peak pitching moment coefficient and drag coefficient. The peak pitching moment coefficient was reduced by 72%, the peak drag coefficient was reduced by 70%, and the lift coefficient hysteresis loop area decreased by 46%. Distributed blowing jet control effectively suppressed the dynamic stall characteristics of the airfoil, making the unsteady load changes gentler.

An experimental investigation on the aerodynamic characteristics and vortex dynamics of a flying wing

AbstractIn this paper, we present a detailed experimental investigation mainly on the vortical flow fields and the associated vortex breakdown phenomena over a non-slender flying wing (sweep angle, ${\rm{\Lambda }}$ = 53°). In the process, the aerodynamic coefficients were also determined using a six-component force balance. Surface oil flow visualisation, surface pressure measurements and particle image velocimetry (PIV) measurements, in various crossflow planes and in a longitudinal plane passing through the leading-edge vortex core, were carried out at various Reynolds numbers to understand the flow field over the non-slender flying wing. Aerodynamic characteristics of the flying wing show local peaks and valleys in the pitching moment coefficient. The surface flow visualisation reveals that the nonlinearity of the pitching moment curve is due to the complex nature of vortical flow structures. The flow visualisation also demonstrates the presence of a wave-like surface pattern, and its size is found to reduce with increasing Reynolds numbers. The present PIV measurements confirm that this wave-like surface pattern is associated with vortex breakdown phenomena. These measurements also reveal that the vortex breakdown has not reached the apex of the wing, even at post-stall angle-of-attack. For pre-stall ( $\alpha $ = 20°) flow regimes, it is observed that the location of the vortex breakdown moves downstream as the Reynolds number increases, but this influence is minimised at near-stall ( $\alpha $ = 25°) and post-stall ( $\alpha $ = 30°) flow regimes. Reconstructed velocity field using the first 10 dominant proper orthogonal decomposition (POD) modes reveals that the nature of the vortex breakdown over the flying wing is a spiral-type vortex breakdown.

Numerical Study on Aerodynamic Characteristics of Tail-stabilized Projectile with Asymmetrical Diversion Groove

A surface diversion groove with a specific geometry and position can influence the laminar flow characteristics of a projectile, which may affect the flight trajectory of an aircraft. The asymmetric flow field around the projectile can be induced by the diversion groove, which can produce an obvious aerodynamic force and moment at the projectile nose for trajectory correction. This study applied a diversion groove structure to the nose of tail-stabilized projectiles to investigate its impact on the aerodynamic characteristics of the projectile. The mathematical expressions for the aerodynamic force and aerodynamic coefficient were established theoretically. The change in the aerodynamic coefficient as a function of the phase angle of the diversion groove was determined. A parametric simulation was employed to investigate how the diversion groove affects the aerodynamic attributes of the projectile across various Mach numbers and angles of attack. The simulation results are consistent with the variation trends of aerodynamic forces and moments with respect to the phase angle of the diverter groove, as predicted by the static mathematical model. These findings demonstrate that the variation trends of the lift coefficient and pitching moment coefficient with respect to the angle β approximate a cosine function. Meanwhile, the variation trends of the yaw force coefficient and yaw moment coefficient with respect to the angle β approximate a sine function. The tail-stabilized projectile with asymmetrical diversion groove achieved a reduction of 1.2% in drag coefficient compared with that of the canard rudder corrective projectile, while the lift coefficient and pitch moment coefficient were increased by 6.4% and 16%, respectively, in the subsonic regime. The static margin of the projectile ranging from 13% to 16%. This study offers valuable insights for the design of corrective structures with diversion grooves and trajectory control.

Comparison of Machine Learning and Classic Methods on Aerodynamic Modeling and Control Law Design for a Pitching Airfoil

Aerodynamic modeling and control law design methods are crucial foundational technologies that enable efficient maneuvering flight of aircraft. Hence, this study focuses on how to construct the aerodynamic model and design control law using machine learning, as well as their differences from classical methods. To conduct this, a modified NACA0012 airfoil with a tail functioning as an elevator is employed as the geometry model. Through the utilization of the rigid dynamic grid method and overlapping grid technology, the pitch of the airfoil and the deflection of the elevator are efficient to execute. Firstly, the pitching moment coefficient is sampled through both steady and unsteady computations. Then, multivariate nonlinear regression (MNR) models and artificial neural network (ANN) models are established based on steady and unsteady sampling data, respectively. Additionally, the obtained model is evaluated using open‐loop control laws. Based on the evaluation, the proportional‐integral‐derivative (PID) control algorithm is used to design the airfoil pitching control law using the MNR model. Meanwhile, deep reinforcement learning (DRL) is used to design the pitching control law using the ANN model. Finally, the PID and DRL controllers are implemented in a CFD environment for airfoil pitching control to verify their effectiveness in application scenarios. The results suggest that models based on both steady and unsteady data can reflect dynamic aerodynamic characteristics. However, using unsteady computation for data sampling significantly reduces time consumption compared to steady computation. Furthermore, the model constructed by ANN may have unexpected excellent characteristics. Both the PID and DRL control laws, designed based on the models, perfectly complete the control process in the CFD environment. This study provides valuable insights for the implementation of controllable maneuvering flight in aircraft.

CFD analysis on the aerodynamic coefficients of an ejection seat system at supersonic speed

In the present study, a 3-D study of the aerodynamic coefficients of an ejection seat system is performed. The analysis is performed using the Reynolds Averaged Navier-Stokes equations where an unstructured grid of the polyhedral cells is used. The aerodynamic coefficients are calculated using a density-based solver and the standard k-ε turbulence model. The investigation is performed at Mach number (Ma) = 1.25 by varying the angle of attack (α) from -15° to 15° with the increment of 5° where the yaw angle (β) is fixed at 0°. According to the findings, the magnitude of the axial force coefficient (CX) increases as α decreases whereas the value of the normal force coefficient (CZ) decreases with the increase of α. Similarly, the value of the pitching moment coefficient (Cm) decreases with the increase of α.

Research of Unsteady Aerodynamic Characteristics of Electrically Controlled Rotor Airfoils with Trailing-Edge Flaps

An electrically controlled rotor (ECR), also known as a swashplateless rotor, eliminates the swashplate system to implement the primary control via the trailing-edge flaps (TEFs), which can result in enhancements in rotor performance, as well as substantial reductions in weight, drag, and cost. In this paper, the unsteady aerodynamic characteristics of the airfoil with TEF of a sample ECR under unsteady freestream condition are investigated. The CFD results are obtained with sliding and overset grid techniques that simulate the airfoil pitching and flap deflection. Comparative analysis of the aerodynamic characteristics under steady and unsteady freestream conditions at different advance ratios is conducted. At various advance ratios, the lift and drag coefficients are higher at a small angle of attack under unsteady freestream condition; however, it is the opposite at a large angle of attack. The peak values of the lift and drag coefficients show an increased trend with the increase in the advance ratio. On the contrary, the pitch moment and flap hinge moment coefficients demonstrate minor variation under unsteady freestream condition. Furthermore, the aerodynamic characteristics of airfoils become more unsteady with variation in the freestream. Therefore, the lift and drag coefficients of the ECR airfoil with TEF show significant differences between steady and unsteady freestream conditions; however, the pitch moment and the flap hinge moment coefficients show little difference.

Wind tunnel testing of a Formula Student vehicle for checking CFD simulation trends

The aerodynamic performance analysis of Formula Student racecars has been mostly done by teams with CFD tools, for time and cost savings, that often lack proper validation. To address this, the FST Lisboa team performed a detailed wind tunnel (WT) test campaign, using a one-third scale model, under different configurations, including variable ride heights, bullhorn appendix, and rear wing flap settings, also replicated in CFD. The simulations used RANS with the SST k-omega turbulence model, with a 13.7 million polyhedral mesh for the test chamber region domain. Both experimental and numerical errors were estimated from the instrumentation and mesh convergence analysis, respectively. Comparisons were made between WT and CFD both in terms of local flow, using tufts for flow visualization, and global flow, using lift, drag, and pitching moment coefficients. Overall, the numerical streamlines agreed very well with the orientations of the tufts in experiments, but some discrepancies were found in regions of cross-flow and high-frequency unsteadiness, mainly caused by limitations of the visualization technique. The gamma transition model in CFD was abandoned as it could not replicate the WT observations. In terms of aerodynamic coefficients, a strong correlation was found between WT and CFD. The parametric studies revealed that the simulations captured the experimental sensitivity to each car setting parameter studied but the uncertainties did not enable a full quantitative evaluation of the aerodynamic performance. The drag reduction system significantly impacted the aerodynamic balance of the racecar, while the current bullhorn design proved to be ineffective. The ride height increase led to higher downforce, mostly due to the higher pitch angle of the vehicle, with negligible variation of the aerodynamic balance. This work validated the team CFD studies, building confidence in that trends estimated in numerical parametric studies are likely to be translated to the real prototype performance.

Experimental studies of airliner aerodynamic characteristics at overcritical angles of attack

PurposeThe main purpose of this study was to determine the basic aerodynamic characteristics of the airliner Tu-154M at the wide range of the overcritical angles of attack and sideslip angles, i.e. α = −900° ÷ 900° and β = −900° ÷ 900°.Design/methodology/approachWind tunnel tests of the Tu-154M aircraft model at the scale 1:20 were performed in a low-speed wind tunnel T-3 by using a six-component internal aerodynamic balance. Several model configurations were also investigated.FindingsThe results of the presented studies showed that at the wide range of the overcritical angles of attack and sideslip angles, i.e. α = −900° ÷ 900° and β = −900° ÷ 900°, the Tu-154M aircraft flap deflection affected the values of the drag and lift coefficients and generally had no major effect on the values of the side force and pitching moment coefficients.Research limitations/implicationsThe model vibration which was the result of flow separation at high angles of attack was the wind tunnel test limitation.Practical implicationsStudies of the airliner aerodynamic characteristics at the wide range of the overcritical angles of attack and sideslip angles allow assessment of the aircraft aerodynamic properties during possible unexpected situations when the passenger aircraft is found to have gone beyond the conventional flight envelope.Social implicationsThere are no social implications of this study to report.Originality/valueThe presented wind tunnel test results of the airliner aerodynamic characteristics at overcritical angles of attack and sideslip angles is an original contribution to the existing not-too-extensive database available in the literature.

Analysis of the aerodynamic characteristics of an ejection seat system using computational fluid dynamics

In the present work, an investigation of the aerodynamic characteristics of an ejection seat occupant is carried out using the commercially available computational fluid dynamics software ANSYS Fluent. 3D Reynolds-averaged Navier–Stokes equations are solved to obtain the aerodynamic coefficients of the ejection seat system. For this analysis, an unstructured grid is generated for the ejection seat occupant using ANSYS meshing. Validation is carried out and the performance of three different turbulence models is analyzed at Mach 0.6. Based on the most suitable turbulence model, further analysis of the aerodynamic coefficients of the ejection seat occupant is calculated at Mach numbers of 0.35, 0.45, 0.55, 0.65, and 0.75. For all values of Mach, the angle of attack is varied from −15° to 15° in 5° increments and the yaw angle is varied from 0° to 60° in 10° increments. Based on the results, it is observed that the magnitude of the axial force decreases with increasing angle of attack and yaw angle. Similarly, the normal force coefficient and pitching moment coefficient decrease with increasing angle of attack. Finally, the side force coefficient, yawing moment, and rolling moment coefficients increase with increasing yaw angle.

Study on the Aerothermoelastic Characteristics of a Body Flap Considering the Nozzle–Jet Interference

A body flap/RCS-integrated configuration is often used to achieve pitch trimming and controlled flight in near space for hypersonic vehicles. Under the high temperature and pressure load induced by the expansion wave at the nozzle exit, the body flap is prone to significant structural deformation, which leads to a change in the resulting moment, even comparable to the control ability, and bring additional challenges to the control system. Based on the CFD/CTD/CSD coupling method, the aerothermoelastic effect on the aerodynamic characteristics and structural deformation of the body flap under jet interaction is systematically studied. Numerical results indicate that the pitching moment coefficients show an increasing trend for all the models, rigid, elastic and thermoelastic, while the increment significantly decreases with the increase in trajectory altitudes. With the increase in deflection angle, the pitching moment coefficients of the three models decrease nonlinearly at high altitude, and the aerothermoelastic effect significantly decreases. At a middle-lower altitudes, the pitching moment coefficient is reversed at a lager deflection angle, and the trailing edge of the body flap presents the deformation characteristics of upward bending, which makes the aerothermoelastic phenomenon degenerate into an aeroelastic problem. From the station along the chord and spanwise direction, the change in displacement increment of the thermoelastic model reflects the competitive relationship between normal stress and thermal stress imposed by jet interaction.