Performance Of Hypersonic Inlet Research Articles

Buzz Characteristics Under Fluid–Structure Interaction of Variable-Geometry Lip for Hypersonic Inlet

Variable-geometry lip structures can effectively broaden the operating range and improve the performance of hypersonic inlets. However, the presence of connection devices leads to a significant reduction in structural stiffness, which can easily trigger fluid–structure interaction (FSI) issues of the lip. In this study, an in-house FSI program is employed to investigate the buzz characteristics under FSI of the lip for a hypersonic inlet. The results indicate that FSI induces instability in the downstream shock train, resulting in the premature occurrence of inlet buzz. The buzz mode induced by FSI significantly differs from the rigid inlet buzz. During the buzz induced by FSI, both violent and mild buzz modes coexist, with these two modes irregularly alternating. During the violent buzz, the inlet achieves a brief start, leading to a fundamentally different evolution of the flowfield structure compared to the rigid model. During the mild buzz, the inlet remains unstart, so the evolution of the flowfield structure is similar to that of the rigid model. In comparison to the rigid inlet buzz, the amplitudes of pressure fluctuations significantly increase, and the dominant frequency decreases for the buzz induced by FSI. Simultaneously, different evolution characteristics of flow frequencies within various inlet regions are observed. Additionally, performance parameters no longer exhibit regular cyclic variations, with significant amplifications in their amplitudes. This study deepens the understanding of the buzz characteristics under FSI of lips for hypersonic inlets and provides a novel perspective on the mechanisms underlying buzz occurrence.

Generalized prediction for self-starting performance of two-dimensional hypersonic inlets

The acceleration self-starting performance of a hypersonic inlet is pivotal for ensuring the stable operation of a scramjet. While the geometric configuration of the internal contraction section (ICS) considerably influences the inlet's self-starting performance, the current prediction model solely considers the internal contraction ratio. To encompass the entire geometric configuration's influence on ICS, a generalized prediction model for the acceleration self-starting Mach number of critical hard unstart two-dimensional hypersonic inlets is proposed. This model calculates the theoretical reattachment pressure rise of the main separation bubble within the actual unstarted flow structure of the hypersonic inlet. Additionally, it computes the theoretical pressure rise assuming the main separation bubble is in a critical state. By comparing these pressures, the model evaluates whether the main separation bubble can be sustained under given incoming flow conditions, predicting the self-starting Mach number. This modeling approach offers broad adaptability to various ICS configurations and incoming flow Reynolds numbers. Each step of the prediction model and the final computational results underwent rigorous evaluation through unsteady numerical simulations. Remarkably, the prediction results demonstrated exceptional alignment with simulation outcomes, surpassing the accuracy of previous prediction methods.

Assessing the Performance of Hypersonic Inlets by Applying a Heat Source with the Throttling Effect

Utilization of a heat source to regulate the shock wave–boundary layer interaction (SWBLI) of hypersonic inlets during throttling was computationally investigated. A plug was installed at the intake isolator’s exit, which caused throttling. The location of the heat source was established by analysing the interaction of the shockwave from the compression ramp and the contact spot of the shockwave with that of the inlet cowl. Shockwave interaction inside the isolator was investigated using steady and transient cases. The present computational work was validated using previous experimental work. The flow distortion (FD) and total pressure recovery (TPR) of the inflows were also studied. We found that varying the size and power of the heat source influenced the shockwaves that originated around it and affected the SWBLI within the isolator. This influenced most of the performance measures. As a result, the TPR increased and the FD decreased when the heat source was applied. Thus, the use of a heat source for flow control was found to influence the performance of hypersonic intakes.

Feasibility of employing the restarting process to evaluate the self-starting ability for hypersonic inlets

Accurate evaluation of hypersonic inlet self-starting performance is very crucial for the stable and efficient operation of scramjet engines. However, few ground hypersonic wind tunnels could provide a continuously and slowly accelerating free stream conditions to directly validate the self-starting ability of hypersonic inlets so far. Some experimental restarting approaches have been explored to evaluate the self-starting performance of hypersonic inlets. To clarify the feasibility of such approaches, the self-starting process and the restarting process caused by withdrawing plug at the exit of a two-dimensional hypersonic inlet were studied using unsteady numerical simulation in this paper. The results show that, even though the freestream Mach number (5.54) of the restarting process is lower than the self-starting Mach number (5.715), the hypersonic inlet can finally rebuild the started flow field owing to the significantly higher moving speed downstream across the inlet internal contraction section caused by the big buzz. Hence, the additional unsteady effect should be suppressed as far as possible during the initial pre-blockage process in the restarting experiment applied to evaluate the self-starting performance of hypersonic inlets.

Integration of inward-turning inlet with airframe based on dual-waverider concept

This paper proposes an inward-turning-inlet–forebody integration methodology based on the dual-waverider concept. For the design concept and the designed flight conditions, both the airframe and inlet benefit from the waverider's high lift-to-drag ratio and the ideal compression performance of its propulsion system. The critical issue is to construct a complex three-dimensional shock wave that accounts for both the inlet and airframe. A method is then proposed for connecting the axisymmetric flow fields of the internal and external flows. The compression surface of the integrated vehicle is then generated from the streamlines obtained with a series of axisymmetric basic flows. Finally, the integration configuration is verified using computational methods, and its viscous effects are analyzed under both on- and off-design conditions via numerical simulations. The results show that a configuration designed using the proposed approach has a high corresponding lift-to-drag ratio and high hypersonic inlet performance.

Effects of the aerothermoelastic deformation on the performance of the three-dimensional hypersonic inlet

The hypersonic inlet is more prone to deform when simultaneously subjected to aerodynamic load and harsh aerothermodynamic load. Moreover, the flow field of the hypersonic inlet is sensitive to configuration. Therefore, it is necessary to investigate the effects of the aerothermoelastic deformation on the flow structure and the performance of the hypersonic inlet. This study develops a loose coupling static aerothermoelastic analysis framework based on the CFD/CSD coupling method, and the one-way and the two-way aerothermal-aeroelastic coupling are both used in the analysis. Furthermore, the effects of the aerothermoelastic deformation on the flow structure and the performance of a three-dimensional hypersonic inlet are studied in detail. The reliabilities of the CFD method and the CFD/CSD coupling method are verified by the validation cases of DLR hypersonic inlet experimental model and the HIRENASD experimental model. The results obtained by the coupling methods are similar. However, the aerothermoelastic deformation obtained through the two-way coupling method is relatively larger, and the effects of the deformation on the inlet performance are more obvious. The maximum of the aerothermoelastic deformation exists at the leading edge of the inlet lip. The deformation changes the shock wave structure near the lip, strengthens the shock wave intensity inside the inlet, increases the length of the separated region and the temperature of the external wall, and changes the flow field of the exit. The aerothermoelastic deformation will lead to the increasing of the mass flow coefficient and the pressure rise ratio; however, it will decrease the total pressure recovery coefficient.

Effects of aeroelasticity on the performance of hypersonic inlet

Effects of the aeroelasticity on the performance of the hypersonic inlet have been investigated numerically in this study. The aeroelasticity has been simulated using the coupled computational fluid dynamics/computational structural dynamics method, which is solved by the in-house code. The unsteady Reynolds-averaged Navier–Stokes equations have been solved in the computational fluid dynamics simulation, and the modal method has been adopted in the computational structural dynamics simulation. Two cases have been utilized to validate the numerical method. Finally, the aeroelasticity has been simulated for inlet plate with different thicknesses. The effects of aeroelasticity on performance parameters and flow structure have been discussed in detail. The results show that the generalized displacements present the “beat” phenomenon in the time domain. The power spectral density of the generalized displacements implies that the aeroelastic instability is mainly caused by the coupling between the fourth- and fifth-order modes. The time-average flow rate coefficient and pressure rise ratio increase relative to the initial value, while the total pressure recovery coefficient decreases. The fluctuation amplitude of the flow rate coefficient is small, while that of the total pressure recovery coefficient and pressure rise ratio are relatively large. Besides, the phases of the three performance parameters are greatly different. Furthermore, the aeroelasticity has significant effect on the shock wave structure especially at the exit of the inlet.

Variable geometry cowl sidewall for improving rectangular hypersonic inlet performance

A variable geometry cowl sidewall control technique is implemented in a generic rectangular hypersonic inlet in this study, in effort to improve inlet performance under both on-design and off-design conditions. Wind tunnel tests and numerical simulations were conducted to investigate the feasibility of the control mechanism, as well as to acquire corresponding inlet flow details. Steady-state pressure distribution was recorded and schlieren photos were obtained simultaneously in the experiment to thoroughly evaluate inlet performance. Combined numerical-experimental results indicate that the variable geometry technique generally enhances inlet performance by two mechanisms. First, the increased cowl sidewall sweepback (CSS) expands the inlet flow spillage window, thus effectively restarting the inlet from an unstart status. The large sweepback cowl sidewall decreases the transverse pressure gradient of the started inlet in the cowl entrance region, compared to an unstart inlet. Second, for a started inlet, decreasing CSS significantly raises mass flux, while slightly lowering total pressure recovery. Smaller CSS, which is preferable for inlet performance improvement, primarily suppresses spillage on the edge of the inlet ramp around the cowl entrance.

Control of Shock/Boundary-Layer Interaction for Hypersonic Inlets by Highly Swept Microramps

The performance of hypersonic inlets is significantly affected by the presence of shock/boundary-layer interactions. To examine the potential of microramps for shock/boundary-layer interaction control in a finite-width duct, a detailed experimental and computational study has been conducted with a separated oblique shock/boundary-layer interaction generated by a 12 deg shock generator at Mach 3.5. Results show that the shock/boundary-layer interaction in the finite-width duct generates complex three-dimensional flow structures with significant swirling nature, and the traditional microramps cannot suppress the separation effectively. Therefore, a type of highly swept microramps with a large chord ratio and small incidence angle is brought forward and investigated. By the precompression effect, the dividing effect, the obstructing effect, and the energizing effect, the highly swept microramps with a height of 0.24 times the boundary-layer thickness show good control capability on the shock/boundary-layer interaction. In addition, the efficiency of the control method for different shock impingement positions is obtained, indicating that the separation can be well controlled when the shock impinges on the aft part of the highly swept microramps.

Numerical Simulation on Air Mass Capture Control of Hypersonic Inlet by Laser Energy

Mechanism of hypersonic inlet performance promoting by “virtual cowl” induced by laser energy is introduced, and the physical model of interaction between laser energy and shocks in hypersonic flow field is established. Through comparing results in this paper with the work performed by Macheret et al.(Princeton University) and analyzing effects of laser energy deposited in hypersonic flow, numerical program and the feasibility of model are validated. At Mach 6, with no laser energy addition, air mass capture ratioKmis 71.87% in this paper, whileKmis 73.17% in the reference. Energy addition module can simulate the formation of shocks induced by laser energy, and the interaction process between laser energy and compression ramp shocks properly. Curves of mass capture ratioKmagrees well with that of the reference.Kmis a little higher than data in the reference, but the variable trend is consistent with each other when laser power is 1kW. When laser power is 0.5kW,Kmvaries a little, and the peak value is 77.36%, which only increases 7.64%, while the value is 7.28% in the reference.

Comparison research of two different improved methods in the 2D hypersonic inlet’s design

The article mainly employed the numerical simulation method to conduct the 2D hypersonic inlet performance research. In this study, the numerical simulation method that includes a two-equation turbulence model was validated through the test case firstly. For the purpose of oppressing unstart phenomena, extending operating range and improving comprehensive performance of the inlet, the bleeding system and isentropic expansion arc surface were applied to the inlet, respectively. Different kinds of configurations were designed through changing the length of the bleeding or changing the radius of the arc surface. The influence of the bleeding and isentropic expansion arc surface on the 2D hypersonic inlet was analyzed systematically based on different objective parameters. The calculative results show that the bleeding system with appropriate value of the length can improve comprehensive performance of hypersonic inlet remarkably either on the design point or the off-design points, though it has a very small negative impact on the inlet in terms of the mass flow ratio coefficient. As for isentropic expansion arc surface, it can improve the inlet’s performance at the design point on the conditions that the radius of isentropic expansion arc surface has an appropriate value. But it has much more disadvantages over the basic configuration reversely when it comes to the off-design conditions or the large size of the radius of isentropic expansion arc surface.

Hypersonic inlet control with pulse periodic energy addition

The theory of pulse periodic energy addition to hypersonic airflow off a vehicle to increase air mass capture and improve the performance of hypersonic inlets at Mach numbers below the design value is explored. The unsteady numerical simulation of hypersonic inlets with energy addition was performed, and pulse periodic energy addition was introduced to discuss the possibility of the steady performance improvement of hypersonic inlets. The numerical simulation of hypersonic inlets at different pulse periodic energy addition rates was carried out. The simulation result shows that the performance parameters are improved at pulse periodic than at fixed values for the same averaged energy addition, and the energy addition rate needed is less at pulse periodic than at fixed values for the same performance parameters of hypersonic inlets.

Mach 4 Performance of Hypersonic Inlet with Rectangular-to-Elliptical Shape Transition

Wind-tunnel testing of a hypersonic inlet with rectangular-to-elliptical shape transition has been conducted at Mach 4.0. This fixed geometry inlet had a geometric contraction ratio of 4.8 and was designed using a quasi-streamline tracing technique to have a design point of Mach 5.7. These tests were performed to investigate the starting and backpressure limits of the inlet at conditions well below its design point. Results showed that the inlet required side spillage holes in order to self-start at Mach 4.0. Once started, the inlet generated a compression ratio of 12.6, captured almost 80% of available air and withstood a backpressure ratio of 30.3 relative to tunnel static pressure. The spillage penalty for self-starting was estimated to be 3.4% of available air. These experimental results, along with previous experimental results at Mach 6.2, indicate that fixed-geometry inlets with rectangular-to-elliptical shape transition are a viable configuration for airframe-integrated scramjets that operate over a significant Mach-number range.

Hypersonic inlet performance from direct force measurements

The problem of establishing one-dimensional values of supersonic combustion inlet performance is approached from the standpoint of direct internal drag measurements. Analytically, one-dimensional inlet component efficiencies can be obtained with mass, momentum, and energy considerations as a function of inlet internal drag. Such efficiency values reflect combined shock wave and viscous losses. Direct measurements of model internal drag were made in a hypersonic shock tunnel to determine the experimental feasibility of this concept for an axisymmetric inlet designed for unity mass flow ratio. With this model, effective internal drag isolation was accomplished by physical separation of the model internal surfaces from adjacent parts of the inlet. With such a technique, a correction must be made to the measured internal drag to account for balance cavity forces. The results of this experiment indicate that, through careful model design, the magnitude of such corrections can be minimized and applied to the measured internal force in an accurate and consistent manner.