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Experimental and numerical studies are carried out to validate the potential of opposing Plasma Synthetic Jet (PSJ) for drag reduction for a hemisphere. Firstly, flow field changes of opposing PSJ are analyzed by comparing the experimental schlieren images and simulation results in a supersonic free stream of Mach number 3. As PSJ is a kind of unsteady pulsed jet, the shock standoff distance increases initially and then decreases under the control of PSJ, which corresponds to the change of the strength of PSJ. Accordingly, the amount of drag reduction of the hemisphere increases initially and then decreases. It is found that there is a short period of “drag rise” during the formation of PSJ before the drag reduction, which is induced by the generation of normal shock waves and the area difference of the cavity wall of PSJ Actuator (PSJA). Secondly, the effects of five parameters, including exit diameter, discharge energy of PSJA, Mach number, static pressure of incoming flow and angle of attack, on drag reduction of opposing PSJ were studied in detail by using numerical method. It is found that the Maximum Pressure Ratio (MPR) has a significant impact on the average drag reduction for a configuration-determined PSJA. For the configuration selected in this study, the flow field of opposing PSJ shows typical Short Penetration Mode (SPM) in a control cycle of PSJ when the MPR is less than 0.89. However, the flow field shows typical Long Penetration Mode (LPM) at some time when the MPR is bigger than 0.89. Relatively better drag reduction is achieved in this case.

The present study proposes a segmented cooling-stream injection structure based on a certain coolant mass flow rate, and numerically investigates the effect of segmented cooling-stream injection on supersonic film cooling. The results indicate that without shock-wave impingement and with helium as the coolant, segmented cooling-stream injection can reduce the mixing between the mainstream and the cooling stream to produce better cooling performance than single injection, especially at larger coolant Mach numbers. However, with nitrogen as the coolant, the cooling effect of the segmented-injection system is very close to that of the single-injection system. Mixing at the impinging region is enhanced significantly when there is an incident shock wave. When the shock wave impinges between the two coolant inlets, segmented cooling-stream injection improves film cooling effectiveness in the midstream and downstream regions more than single injection because only part of the cooling stream undergoes the enhanced mixing effect of the shock wave. The advantage of segmented injection is reduced when the impinging region is behind the second coolant inlet. The further downstream the impinging region, the smaller the associated advantage.

The energy efficiency of a supersonic ramjet engine with detonation combustion of a hydrogen-air mixture in an axisymmetric Laval nozzle is considered. First of all, within the framework of the one-dimensional (1D) theory, it is shown that extreme combustion corresponding to the Chapman-Jouguet (C.-J.) deflagration, like constant pressure combustion, surpasses detonation in energy efficiency. However, this advantage becomes insignificant with an increasing the free-flow Mach number. Based on the results of the 1D theory, an axisymmetric Laval nozzle is constructed, which provides steady-state detonation combustion of a stoichiometric hydrogen-air mixture at an altitude of about 40 km and with freestream Mach numbers 9 and 12. The efficiency of detonation combustion in this case is about 35% instead of the maximum 65% according to the 1D theory.

A novel method for close-proximity formation flying under differential atmospheric drag using Coulomb forces is investigated for applications in Earth sensing, space-situational awareness (SSA), and aeronomy. Objects in LEO are supersonic with respect to the ambient environment, creating a thinned out wake region behind the craft as it travels through the ionosphere. Objects within this wake experience little drag acceleration and are able to attain voltages much greater than in the ambient ionospheric plasma, creating implications for the design and control of close-proximity leader–follower spacecraft pairs. The proposed system consists of a leader craft with a set of affixed, conducting spheres and a charged follower craft located in the wake of the leader. The differential drag acceleration between the leader and follower craft is countered by a controlled Coulomb repulsion to maintain precise separation. The charged structure on the rear of the leader craft is designed such that the charged follower craft sits in an electrostatic potential well which opposes off-axis perturbations. A conceptual method for controlling such a pair without the use of propellant using a set of charged spheres is investigated, with nonlinear models of the system’s relative motion derived and discussed. Linearized models are used to demonstrate the local controllability of the system to demonstrate the proposed system’s merit. This linear analysis is used to derive conditions on controllability and control performance under different charge geometries and environmental assumptions.

The accurate prediction of the sonic boom is essential for the development of a future low-boom supersonic aircraft. In this paper, the best practice for predicting near-field pressure signatures w...

Predictions using the launch ascent and vehicle aerodynamics (LAVA) solver framework for the NASA shock–plume interaction biconvex model and C608 supersonic aircraft prototype geometries are compar...

Supersonic flows around parachute two-body systems are numerically investigated by solving the compressible Navier-Stokes equations. In the present study, both rigid and flexible parachute models are considered, which comprise a capsule and a canopy. The objective of the present study is to investigate the effects of the Martian atmosphere on the unsteady flows produced by these parachute two-body models and the structural behavior of the flexible canopy. It was found that in the Martian atmosphere, the supersonic rigid parachutes with shorter trailing distances exhibited weaker aerodynamic interactions between the capsule wake and canopy shock, resulting in a smaller pressure distribution on the typical surfaces of the canopy. By contrast, because the flow modes around the flexible parachute in the Martian atmosphere were similar to those of the rigid parachute under the air conditions of the wind tunnel tests, the canopy shape was almost unchanged. When a new canopy material was designed by decreasing the Young’s modulus and damping coefficient, an area oscillation phenomenon was observed in the flexible parachute with a trailing distance of 10 in the Martian atmosphere. Consequently, the Martian atmosphere (low density and pressure) has a significant effect on the aerodynamic performance of the flexible parachute system.

Current design constraints have encouraged the studies of aeroacoustics fields around compressible jet flows. The present work addresses the numerical study of subgrid scale modeling for unsteady turbulent jet flows as a preliminary step for future aeroacoustic analyses of main engine rocket plumes. An in-house large eddy simulation (LES) tool is developed in order to reproduce high fidelity results of compressible jet flows. In the present study, perfectly expanded jets are considered because the authors want to emphasize the effects of the jet mixing phenomena. The large eddy simulation formulation is written using the finite difference approach, with an explicit time integration and using a second order spatial discretization. The energy equation is carefully discretized in order to model the energy equation of the filtered Navier-Stokes formulation. The classical Smagorinsky model, the dynamic Smagorinsky model and the Vreman models are the chosen subgrid scale closures for the present work. Numerical simulations of perfectly expanded jets are performed and compared with the literature in order to validate and compare the performance of each subgrid closure in the solver., Comment: conference paper - AIAA Aviation 2016. arXiv admin note: substantial text overlap with arXiv:2212.12353, arXiv:2212.12365

The variable geometry supersonic inlet tends to decrease the throat area to reduce the Mach number upstream of the terminal shock, so as to reduce the flow loss. However, excessive Internal Contraction Ratio (ICR) exposes the inlet to a greater risk of unstart, which inevitably results in a process of increasing the throat area to aid the inlet restart. In the above throat regulation process, the inlet undergoes the start, unstart, and restart states in turn. In order to reveal the flow structure and mechanism of this process, a two-dimensional unsteady numerical simulation combined with a dynamic mesh technique were employed. The shock-on-lip Mach number of the studied inlet is 4.0 and the flight angle of attack is + 6°. Analysis was focused on the state with a freestream Mach number of 3.0. The results clearly show that the flow response hysteresis appears, and restart is only realized when the throat area is obviously increased as compared to that of unstart due to the historical unstart flow structure. In addition, three typical flow fields were analyzed, and it is found that the separation ahead of the inlet was the key factor affecting the hysteresis. Finally, unstart and restart boundaries of the inlet were discussed, and the factors influencing its deviation from the typical boundaries of dual-solution area were analyzed. The newly predicted unstart and restart boundaries are much closer to the CFD results.

We conduct large-eddy simulations to investigate the role of thermal nonuniformity on the development of instability modes in the shear layer of supersonic M=1.5 jets. We develop an arbitrary high-...

In this paper, the low-speed aerodynamic characteristics of a basic wing–fuselage supersonic biplane incorporating trailing-edge flaps are assessed using both wake and balance measurements. The eff...

Experimental investigations of two interacting spheres were carried out in a Mach 4 flow with a static pressure of 2.666 Pa. This study was performed in the MARHy wind tunnel, an adjustable supersonic or hypersonic low-density facility. The main purpose is to estimate how the interactions between spherical debris modify the aerodynamic forces during their atmospheric entry. Indeed, around 80 km, altitude experimentally reproduced thanks to MARHy flow conditions, most debris fragment. The multiple debris, resulting from fragmentation, interact with each other’s, which may impact their trajectory. The purpose of this work is to analyze the behavior of a moving sphere placed in the wake of another one. The moving sphere is suspended with two thin wires, so it can almost freely move in the flow direction and physically respond to the acting flow forces of the wake of the first one. Aerodynamic forces are then calculated from the sphere angle deflection. The qualitative analysis of the experimental results will be discriminated first with respect to the trajectory of the second sphere, and secondly with respect to the evolution of the aerodynamic forces according to the sphere.

Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow, which always accompanies aerodynamic performance penalties. A loss reduction method for smearing the passage shock foot via Shock Control Bump (SCB) located on transonic compressor rotor blade suction side is implemented to shrink the region of boundary layer separation. The curved windward section of SCB with constant adverse pressure gradient is constructed ahead of passage shock-impingement point at design rotor speed of Rotor 37 to get the improved model. Numerical investigations on both two models have been conducted employing Reynolds-Averaged Navier-Stokes (RANS) method to reveal flow physics of SCB. Comparisons and analyses on simulation results have also been carried out, showing that passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot for moderate adverse pressure gradient as well as suppression of boundary layer separations and secondary flow of low-momentum fluid within boundary layer. It is found that adiabatic efficiency and total pressure ratio of improved blade exceeds those of baseline at 95%-100% design rotor speed, and then slightly worsens with decrease of rotatory speed till both equal below 60% rated speed. The investigated conclusion implies a potential promise for future practical applications of SCB in both transonic and supersonic compressors.

This paper examines various methods for phase averaging the flowfield of a sweeping supersonic jet emanating from a scaled fluidic oscillator. Determination of the phase-averaged flowfield is chall...

Supersonic boundary-layer transitions triggered by a diamond-shaped roughness element were investigated by direct numerical simulation based on a finite volume formulation with a high-order minimum...

Suppression of shock-induced separation on all the walls of a rectangular duct using arrays of microvortex generators was experimentally investigated. The flowfield was analyzed using normal and in...

The discontinuity at the thermal protection system tiles covered on an aerospace vehicle is modeled as an open cavity to numerically investigate rarefied hypersonic flows over it under various flow control styles using the direct simulation Monte Carlo method. The influence of flow control on flow structures (e.g., shock wave and vortex) and aerothermal properties (e.g., temperature distribution and wall heat flux) is analyzed in detail. The flow structures and aerothermal properties of the cavity flowfield are both changed greatly by the square control block on the cavity opening line, such as a reduction of about 79.5% of the maximum of wall heat fluxes near the reattachment corner, while these changes are gradually weakened with the control block moving down. The control blocks with different shapes still have the ability to change flow structures and thus the heat flux coefficients over the cavity surfaces, but shape has little effect, and the square control block has the strongest wall heat flux compared with the circular, triangular, and wedged control blocks. For the supersonic jet with different directions, the leftward jet has the best performance in eliminating high temperature gas inside the cavity and over the downstream wall and reducing wall heat fluxes over cavity surfaces, followed by the downward jet. The main reason is that the leftward and downward jets can effectively prevent momentum and energy exchanges between freestream with cavity surfaces.

A study is undertaken to assess the dynamic control's effectiveness at two supersonic Mach numbers, 2.2 and 2.58 being considerably large Mach number being very significant in the design of High-speed aerospace vehicles. The relief available to the flow while the shear layer exits from the nozzle and enters the enlarged duct. The investigation area ratio is 4.84, and the NPR tested is from 3 to 11, covering the various level of expansion. The investigation was done for eight sizes of circular duct cases. Still, the results are discussed only when a significant change in the wall pressure is noticed during the experimentation. The dynamic control is located at a 6.5 mm radial distance from the primary jet's axis, with a 1 mm diameter generating at sonic Mach from the active flow control management. For lower duct length at NPRs, the wall pressure assumes large values. Later, with a rise in the duct section and NPR, a progressive decline in the pressure values is seen due to the decrease in adverse pressure levels and the intensity of the oblique shock waves. The duct flow mild oscillations are seen for duct section 3D, 5D, 6D, and 10D. The duct length portrays a crucial role while determining the stream development in the enlarged circular tube. When the duct's size is 6D and 10D, the low wall pressure is seen. For NPR = 5, the wall pressure takes moderate values due to the influence of the tube. At NPR = 7, 9, and 11, the pressure rise is limited within the reattachment zone, and ambient conditions are nearly attained within a short duct length. Despite nozzles, they are over-expanded flow control effects in a rise of pressure at Mach 2.58 for duct size 6D and 10D. For the duct size 10D, the attainment of ambient pressure is further advanced towards the upstream.

Compressible cavity flows exist widely in aeronautical vehicles. The cost of the computational fluid dynamics to solve compressible cavity flows is high. To address this issue, in this work, a nove...

Selective metallization on glass surface technique has lot of potential applications in optoelectronics. Particularly, ultra-fine copper lines provide better optical transparency, which may enable to meet the demands of novel optoelectronic devices. In this paper, we propose a new manufacturing procedure to overcome the limitations of the traditional process and to achieve narrower copper lines on the glass surface. We, first, irradiate the glass surface by the femtosecond laser for more precise ablation, and then deposit the silver nitrate particles onto the irradiated region by using a supersonic particle deposition system. Copper lines with a linewidth of 1 µm were achieved by means of this new procedure, which is, to best our knowledge, the narrowest linewidth ever reported so far using laser direct writing and the subsequent electroless copper plating. We also demonstrate a highly transparent heater (i.e., 97% optical transparency) as an application of the proposed manufacturing approach.

With the increasing altitude and speed of the hypersonic vehicles, the temperature of the combustion chamber wall must be substantially decreased to ensure thermal and structural reliabilities. As a great supplement to the main regenerative cooling method without additional coolant, film cooling with onboard hydrocarbon fuel is widely used to decrease the wall heat flux. In this paper, a typical cylindrical film injection hole is employed to study the flow and heat transfer characteristics of gaseous film cooling with hydrocarbon fuel in the supersonic main flow, and the effects of transverse-non-uniform distributions of blowing ratio and injection angle are analyzed in detail. The results show that entrainment process between cooling films with different blowing ratios and injection angles would influence the film cooing efficiency to different degree, and proper combination of these two factors can improve the film cooling efficiency and decrease the flow rate of the cooling film. And the empirical correlation of supersonic hydrocarbon film cooling efficiency is established for the uniform blowing ratio cases. For the transverse non-uniform blowing ratio, the overall cooling efficiency decreases due to the spanwise and height-wise ejections between the two groups of cooling film. And for the cases with combination of transversely non-uniform blowing ratio and unequal injection angle, the film cooling effectiveness per blowing ratio would increase by 8.0%.

Aircraft infrared (IR) signature studies are complex, due to their dependence on several parameters, e.g., line of sight (LOS) and viewing aspect (LOS^). There is no equivalent counterpart of the w...

The solid-fueled Scramjet is an interesting option for supersonic combustion ramjet. It shows significant advantages such as simple fuel supply and compactness, avoiding the complex system of tanks and pipelines that encountered in liquid-fueled Scramjets. The solid-fueled Scramjet could be the simplest air-breathing engine for the hypersonic flight regime. This paper presents a comprehensive and systematic review of the research progress on solid-fueled Scramjet in various institutes and universities. It summarizes a progress overview of three types of the solid-fueled Scramjet, which covers a wealth of landmark numerical and experimental results. Based on this, several relevant key technologies are proposed. Several inherent scientific issues are refined, such as the mixing mechanism of multi-phase flow and supersonic airflow, ignition and combustion mechanism of the condensed phase in a supersonic airflow, and coupling mechanism of gas and solid phase in a supersonic flow. Finally, the historical development trend is clarified, and some recommendations are provided for future solid-fueled Scramjet.

In this paper, the effects of non-isothermal oxidation on transient conjugate heat transfer of the cryo-supersonic air-quenching are investigated based on a double-layered oxidation kinetics model, while a unified conjugate heat transfer formula is developed to synthetically consider the near-wall turbulence, non-isothermal oxidation, and surface radiation. The comparison between numerical and experimental results are also presented to check the validity of the developed model. The results indicate that the film growth has some degree of inhibition to the conjugate heat transfer, in particular, the stagnation temperature increases linearly by about 5 K per 100 μm increase in film thickness, and the effective conjugate heat transfer coefficient in the stagnation region decreases linearly by about 55 Wm-2K-1 per 100 μm increase in film thickness. Moreover, the oxide film would have little impact on transient conjugate heat transfer when the near-wall velocity is higher due to the effect of viscous dissipation.

Nonintrusive laser diagnostics are used for the measurements of metastable nitrogen molecules in the lowest excited electronic state, N2(A3Σu+), in a nonequilibrium flow, blowdown supersonic wind t...