Unstart dynamics of hypersonic Busemann inlet at non-zero angles of attack

The propensity of a streamtraced hypersonic Busemann or “sugar scoop” inlet to unstart is experimentally and computationally investigated at positive and negative angles of attack. The backpressure was applied at the isolator exit in the form of an aerodynamic blockage from a high-momentum jet to trigger inlet unstart. High-speed schlieren photography revealed two distinguishable responses during backpressured operation at different angles of attack. At low backpressures, a “weak response” was observed where pseudoshocks were disgorged from the inlet’s internal contraction passage but with a supersonic flow channel present at the cowl closure. When the backpressure was increased, a “strong response” was observed where the entirety of the flow entering the internal contraction segment of the inlet was subsonic, which evidenced the completion of the unstart process. The weak response was observed at a lower backpressure setting with increasing angles of attack. However, a smaller subsequent increase in backpressure was required to achieve the strong response at lower angles of attack compared to higher angles of attack. Computational simulations anchored on the schlieren imagery and two-dimensional pressure fields, measured in the inlet external contraction region, were performed to probe the disparity in the unstart response across different angles of attack. The differences in operational limits and sensitivity margins were attributed to the viscous effects and momentum fluxes through the inlet/isolator.

In the present study, multiple-fan active control wind tunnel tests are conducted to investigate the aerodynamic forces on a 5:1 rectangular cylinder in sinusoidal streamwise winds with different angles of attack (AoA). The effects of the frequency, amplitude, and AoA of the sinusoidal flow on the statistical parameters, spectral characteristics, and spanwise distributions of drag, lift, and moment coefficients are analyzed. Results show that each force has two components: the one induced by the approaching velocity oscillation and the one induced by wake vortex-shedding—this is quite different from that in the smooth flows, where the fluctuating forces are totally due to wake vortex-shedding. For each force, changes of the two components and their relationship with the frequency, amplitude, and AoA are presented. The drag fluctuations are generally dominated by the approaching velocity oscillations, whereas the lift and moment are more sensitive to wake vortex-shedding. Therefore, the drag force has better spanwise correlations than the lift and moment forces. Meanwhile, at a non-zero AoA, the inflow amplitude has different effects on the vortex-shedding-induced component as that at a zero AoA. The differences of spanwise distributions between the sinusoidal flow cases and the smooth flow cases are analyzed.

A study is made of the flow of a compressible gas in a laminar boundary layer on swept-back wings of infinite span in a supersonic gas flow at different angles of attack. The surface is assumed to be either impermeable or that gas is blown or sucked through it. For this flow and an axisymmetric flow an analytic solution to the problem is obtained in the first approximation of an integral method of successive approximation. For large values of the blowing or suction parameters, asymptotic solutions are found for the boundary layer equations. Some results of numerical solution of the problem obtained by the finite-difference method are given for wings of various shapes in a wide range of angles characterizing the amount by which the wings are swept back and also the blowing or suction parameters. A numerical solution is obtained for the equations of the three-dimensional mixing layer formed in the case of strong blowing of gas from the surface of the body. The analytic and numerical solutions are compared and the regions of applicability of the analytic expressions are estimated. On the basis of the solutions obtained in the present paper and studies of other authors a formula is proposed for the calculation of the heat fluxes to a perfectly catalytic surface of swept-back wings in a supersonic flow of dissociated and ionized air at different angles of attack. Flow over swept-back wings at zero angle of attack has been considered earlier (see, for example, [1–4]) in the theory of a laminar boundary layer. In [5], a study was made of flow over swept-back wings at nonzero angle of attack at small and moderate Reynolds numbers in the framework of the theory of a hypersonic viscous shock layer.

In the present study a single expansion ramp nozzle (SERN) system is studied with nozzle pressure ratio (NPR) of 2.9. An interaction of subsonic ambient inflow with M∞ of 0.4 which is induced during take-off of a typical hypersonic vehicle with SERN is considered. Different ambient angles of attack of 0°, −5°, −10° and −15° have been studied. Results obtained by numerical computation are validated by available experimental data. A significant amount of separation is developed at the walls (on both ramp and flap) due to shock wave-boundary layer interaction (SWBLI). Due to the ambient inflow at different angles of attack the separation patterns changes from restricted shock separation (RSS) to free shock separation (FSS). This transition happens when angle of attack changes from 0° to −5° at ambient inflow Mach number of 0.4. Due to ambient stream interaction shock wave oscillates and wall pressure distribution on both ramp and flap changes with time for different ambient inflow conditions. This flow field instability is expressed by the distribution of root means square (RMS) value of pressure over ramp and flap. The main purpose of SERN is generation of thrust of hypersonic vehicles. So thrust and thrust vector angle are important parameters of SERN. Thrust and thrust vector angle of SERN is related to separation pattern. This study gives an insight to the flow structure in SERN when there is a strong ambient inflow and can be very beneficial to the design and performance analysis of SERN.In the present study a single expansion ramp nozzle (SERN) system is studied with nozzle pressure ratio (NPR) of 2.9. An interaction of subsonic ambient inflow with M∞ of 0.4 which is induced during take-off of a typical hypersonic vehicle with SERN is considered. Different ambient angles of attack of 0°, −5°, −10° and −15° have been studied. Results obtained by numerical computation are validated by available experimental data. A significant amount of separation is developed at the walls (on both ramp and flap) due to shock wave-boundary layer interaction (SWBLI). Due to the ambient inflow at different angles of attack the separation patterns changes from restricted shock separation (RSS) to free shock separation (FSS). This transition happens when angle of attack changes from 0° to −5° at ambient inflow Mach number of 0.4. Due to ambient stream interaction shock wave oscillates and wall pressure distribution on both ramp and flap changes with time for different ambient inflow conditions. This flow field...

An experimental study was carried out on an axisymmetric supersonic inlet with external compression in order to investigate the buzz phenomenon at different angles of attack and mass flow rates. The model was equipped with accurate and high-frequency pressure sensors, and the tests were conducted at Mach numbers varying from 1.8 to 2.5, for various angles of attack. Shadowgraph visualization technique, together with a high-speed camera, was used to provide the visual description of the shock structure in front of the inlet and to study the characteristics of buzz. Furthermore, pressure distribution over the spike surface was measured using several pressure sensors. Frequency of the buzz and shock displacement were measured by inspection of visualization pictures in each test. The obtained data from shadowgraphs were compared with those obtained from pressure measurements, and good agreement was found between them. The results revealed that for a moderate value of mass flow rate, the frequency of shock oscillation decreases as Mach number increases. Further, by increasing angle of attack, the shock displacement of oscillation will increase. At non-zero angles of attack, the displacement and frequency of shock motion show different behaviors on the leeward and windward sides of the body.

Boundary layer stability was analyzed for the HIFiRE-1 flight vehicle geometry for ground tests conducted at the CUBRC LENS I hypersonic shock test facility and the Langley Research Center (LaRC) 20- inch Mach 6 Tunnel. Boundary layer stability results were compared to transition onset location obtained from discrete heat transfer measurements from thin film gauges during the CUBRC test and spatially continuous heat transfer measurements from thermal phosphor paint data during the LaRC test. The focus of this analysis was on conditions at non-zero angles of attack as stability analysis has already been performed at zero degrees angle of attack. Also, the transition onset data obtained during flight testing was at nonzero angles of attack, so this analysis could be expanded in the future to include the results of the flight test data. Stability analysis was performed using the 2D parabolized stability software suite STABL (Stability and Transition Analysis for Hypersonic Boundary Layers) developed at the University of Minnesota and the mean flow solutions were computed using the DPLR finite volume Navier-Stokes computational fluid dynamics (CFD) solver. A center line slice of the 3D mean flow solution was used for the stability analysis to incorporate the angle of attack effects while still taking advantage of the 2D STABL software suite. The N-factors at transition onset and the value of Re(sub theta)/M(sub e), commonly used to predict boundary layer transition onset, were compared for all conditions analyzed. Ground test data was analyzed at Mach 7.2 and Mach 6.0 and angles of attack of 1deg, 3deg and 5deg. At these conditions, the flow was found to be second mode dominant for the HIFiRE-1 slender cone geometry. On the leeward side of the vehicle, a strong trend of transition onset location with angle of attack was observed as the boundary layer on the leeward side of the vehicle developed inflection points at streamwise positions on the vehicle that correlated to angle of attack. Inflection points are a strong instability mechanism that lead to rapid breakdown and transition to turbulence. The transition onset location on the windward side of the vehicle displayed no trend with angle of attack or freestream Reynolds number and transition was observed farther down the vehicle than observed on the leeward side of the vehicle. In analysis of both windward and leeward sides of the vehicle, use of the N factor methodology to develop trends to predict boundary layer transition onset showed improvements over the Re(sub theta)/M(sub e) empirical correlation methodology. Stronger correlations and less scatter in the data were observed when using the N factor method for these cases.

本文根据以往翼身融合体飞机的一些设计经验，设计了一种翼身融合体飞机气动外形，并研究分析了翼身融合体飞机初始模型在不同风速和迎角的情况下的气动特性。利用CATIA生成翼身融合体飞机初始模型，由于考虑到该型飞机的具体尺寸、选用翼型以及制作重量，所以飞机的巡航速度设计为30 m/s，通过风洞实验和FLUENT计算流体力学软件计算得出初始翼身融合体飞机模型的速度云图和压力云图以及初始模型飞机周围流场分布情况，然后根据FLUENT计算流体力学软件得出的初始模型流场分布情况，确定4种在初始模型基础上经过改进的模型，分别为涡流器模型、翼梢小翼模型、鸭翼模型、平尾模型，研究经过改进的4个模型气动特性(包括升力系数，阻力系数，升阻比)随速度和迎角的影响的思路是：建立精确的改进涡流器模型、改进翼梢小翼模型、改进鸭翼模型、改进平尾模型，利用CFD计算流体力学软件计算在不同风速迎角分别从−4˚到14˚下飞机的气动特性，并分析在不同速度和迎角下的飞机的气动特性随迎角的变化规律，为今后翼身融合飞机的安全高效飞行提供理论依据。并对改进的4种模型在不同速度和迎角下的升阻比等气动特性与初始模型进行对比，再根据不同的改进模型在不同飞行状态下所具有的气动特性探讨其改进模型的用途及其发展价值。 This article based on the design experience of wing body fusion aircraft before designed a wing body fusion aircraft aerodynamic shape, and analyzed initial model of the wing fusion aircraft with different wind speed and angle of attack. In the process of research, CATIA was used to generate the initial model of wing body fusion aircraft. Taking into account the specific size of the aircraft, the choice of airfoil and for weight, the cruise speed of the aircraft is designed to be 30 m/s. Through the wind tunnel experiment and FLUENT computational fluid dynamics soft-ware calculations, we can get the initial model of wing fusion aircrafts speed cloud and pressure cloud and the distribution of the surrounding flow field. On the basis of FLUENT computational fluid dynamics software and the distribution of the surrounding flow field, four improved models based on the initial model were identified: Eddy swirl model, wingtip winglet model, canyon model, and flat tail model. This paper researched the aerodynamic performance of the four improved wing-body fusion aircraft models at different angles of attack and wind speed (including lift coefficient, drag coefficient, lift drag ratio). The idea of the impact is to establish an accurate improved swirl model, to improve the wing winglet model, to improve the duck wing model, to improve the flat tail model. The research has used CFD to calculate the aerody-namic characteristics of the airplane at different wind of angles at −4˚ to 14˚, and analyzed the aerodynamic characteristics of the aircraft at different speeds and angles of attack. It can pro-vide a theoretical basis for the safe and efficient flight of the aircraft. This research compared the aerodynamic characteristics of the improved four models with the initial model at different speeds and angles of attack. And then it used the aerodynamic characteristics of different im-proved models under different flight states to explore the use of the improved model and its development value.

Turbomachinery rotor blades experience unsteady lift forces due to turbulence and circumferential variation in the inflow velocity. These forces can be a source of unwanted vibration and radiated sound. The unsteady force spectrum is often modelled to first order using a flat plate approach. However, it is known that mean loading effects can have a significant influence on airfoil unsteady lift for sinusoidal gusts with certain wavevector orientations [1]. To this end, the Atassi [1] gust response model for airfoils with camber and non-zero angle of attack has been incorporated into a rotor turbulence ingestion model [2] to determine the effect on the unsteady force spectrum. The Atassi gust model describes the interaction between a two-dimensional sinusoidal gust with a cambered foil placed at a small angle of attack to the mean flow. To apply this single harmonic gust response to the turbulence ingestion model, the gust responses for sinusoidal components with a distribution of wavevector angles are integrated based on an isotropic turbulence energy spectrum. The updated turbulence ingestion model is then applied to a parametric study investigating the effects of foil geometry (specifically camber and non-zero angle of attack) on the unsteady forces produced from turbulence ingestion. The baseline geometry used for this study is the Sevik [3] rotor. To assess the sensitivity of the unsteady force spectrum to camber, predictions for a 0, 6, and 12% cambered foil with zero thickness at 0° angle of attack are generated. For these cases, foil camber effects are found to be largely negligible, except at low frequencies where the additional forces generated by streamwise gust components become dominant. The influence of non-zero angle of attack on the turbulence ingestion rotor force is also relatively minor for small angles of attack, but an increase of a few dB is predicted at higher frequencies as the angle of attack is increased to 15°. These results are generally consistent with measurements reported in the literature, where camber and angle of attack effects are shown to be weak, and the overall effects are much smaller than the changes in the sinusoidal gust responses due to the streamwise symmetry and averaging effect that occurs with isotropic turbulence. While the turbulence ingestion model was successfully extended to account for foil geometry effects using the Atassi gust response model, there is continued interest to explore the effects on the unsteady force spectrum when the simplifying assumption of isotropic turbulence is removed.

Experiments are carried out to study flow and heat transfer characteristics over NACA0018 aerofoil when the body approaches the wall of a wind tunnel. Investigations have been done to study the effect of wall proximity due to flow separation around the body at Reynolds number 2.5 × 105, different height ratios and various angles of attack. The static pressure distribution has been measured on upper and lower surfaces of the aerofoil. The results have been presented in the form of pressure coefficient, drag coefficient for different height ratios. Pressure coefficient values are decreased and then increased on the lower surface of the aerofoil and decreased on the upper surface of the aerofoil at all angles of attack. The negative pressure coefficient and drag coefficient decreases as the body approaches the upper wall of wind tunnel. The maximum value of drag coefficient has been observed at an angle of attack 30° for the aerofoil at all height ratios. The Heat transfer experiments have been carried out under constant heat flux condition. Heat transfer coefficients are determined from the measured wall temperature and ambient temperature and presented in the form of Nusselt number. The variation of local as well as average Nusselt number has been shown with non dimensional distance for different angles of attack and for various height ratios. The local as well as average Nusselt number decreases as the height ratio decreases for all non-dimensional distance and angles of attack respectively. Maximum value of average Nusselt number has been observed at an angle of attack 40°.

Flow structures behind a vibrissa-shaped cylinder at different angles of attack: Complication on vortex-induced vibration

Nature-inspired flying robots are beneficial than other multi-rotor or fixed wing analogs, in many aspects. As wings play the key role on the hovering and maneuvering conditions of flying insects, structural functions and aerodynamic performances of the insect wings are needed to be analyzed for designing more effective wings for insect-sized flying robots. This study describes the method for experimental analysis of aerodynamic and vibration characteristics of dragonfly (Erythemis Simplicicollis) forewings and hindwings. Vibration testing of the dragonfly wings has been conducted to obtain natural frequencies and mode shapes of the wings. The wings have also been examined in a suction wind tunnel having pistol-grip sting balance to illustrate the vibration and aerodynamic characteristics. The structural aerodynamic response of the wing has been determined at different freestream velocities and at different angles of attack. From the experimental results, the deformation response and the coefficients of drag and lift of the insect wings have been obtained for different Reynolds numbers and angles of attack. The coefficient of lift of the wings increases with the Reynolds number and angle of attack. The coefficient of drag of the wings also increases with the Reynolds number and angle of attack.

Conventional propellers operating at negative thrust conditions, even at 0 deg angle of attack, are characterized by flow separation and significantly different noise emissions than at positive thrust conditions. Operating the propeller at nonzero angles of attack at negative thrust conditions can further impact aerodynamic performance and far-field noise emission. This paper studies these effects using lattice-Boltzmann very large eddy simulations coupled with the Ffowcs Williams and Hawkings analogy. At positive thrust, operation at 10 deg angle of attack increases thrust along the freestream direction by approximately 3% compared to operation at 0 deg angle of attack, while efficiency remains constant. Conversely, the negative thrust condition shows approximately a 7% decrease in thrust magnitude and a 10% reduction in regenerated power. In this condition, the positively cambered blade sections exhibit dynamic stall, resulting in broadband fluctuations of up to 10% of the mean loading near the blade tip. The nonzero angle of attack induces opposite variations in absolute blade loading between positive and negative thrust conditions, resulting in opposite changes in the noise directivity. At positive thrust, noise increases in the region from which the propeller is tilted away (i.e., below the propeller at a positive angle of attack), while the opposite occurs at negative thrust. The varying blade loading over the azimuth results in destructive interference between loading and thickness noise for the negative thrust case at the 10 deg angle of attack. These findings highlight the crucial role of considering nonzero angles of attack in propeller design and optimization analyses.

This paper studies the effect of operation at non-zero angles of attack on the aerodynamic performance and far-field noise emissions of an isolated propeller operating at positive and negative thrust conditions. To achieve this, scale-resolved lattice-Boltzmann very large eddy simulations coupled with the Ffowcs Williams & Hawkings analogy have been used. The results show that when the propeller operates with a 10◦ angle of attack at the positive thrust condition, the blade loading increases on the advancing side and decreases on the retreating side, leading to a 9.6% increase in integrated thrust (when computed along the propeller axis) and a negligible increase (0.1%) in propeller efficiency. Conversely, at the negative thrust condition, the operation at 10 deg angle of attack results in a 7.9% decrease in thrust magnitude and an 11.1% reduction in energy-harvesting efficiency. In this condition, the positively cambered blade sections exhibit dynamic stall at the 10◦ angle of attack, resulting in broadband fluctuations of up to 10% of the mean loading. As a result of the opposite change in absolute blade loading in the negative thrust condition compared to the positive thrust condition at the 10◦ angle of attack, the change in the noise directivity is also the opposite. Whereas in the positive thrust case, the noise increases in the region from which the propeller is tilted away (i.e., below the propeller at a positive angle of attack), in the negative thrust case, it is the other way around. This study highlights the need to account for non-zero angles of attack in propeller design and optimization analyses.

An experimental study of the breakdown of leading-edge vortices on diamond, cropped, delta, and double delta wings during dynamic pitching

Numerical study on the supersonic near-wake flow structures of a cylinder at an angle of attack