Journal articles: 'Hypersonic inlet design'

1

Zhai, Jian, Chen-An Zhang, Fa-Min Wang, and Wei-Wei Zhang. "Alleviation of lateral spillage of two-dimensional hypersonic inlet using waverider-configuration chines." International Journal of Modern Physics B 34, no. 14n16 (June 4, 2020): 2040074. http://dx.doi.org/10.1142/s0217979220400743.

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Hypersonic inlet is an important part of the propulsion system of hypersonic air-breathing vehicles. However, the performance of the two-dimensional hypersonic inlet, a major type of hypersonic inlets, is considerably deteriorated for lateral spillage. In this study, waverider-configuration chines mounted on the lateral sides of a two-dimensional three-staged external-compression hypersonic inlet for a Mach number of 6.0 are investigated to determine their ability to alleviate the lateral spillage. The chines are built by using a waverider design method. The numerical results suggest that a severe flow spillage induced by three-dimensional effect shows up near the lateral edge of the inlet without chines, which degrades the mass-flow ratio and flow uniformity. In contrast, the waverider-configuration chines effectively alleviate the lateral spillage. Consequently, the mass-flow ratio and the flow uniformity are both improved significantly.

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Saheby, Eiman B., Huang Guoping, and Anthony Hays. "Design of hypersonic forebody with submerged bump." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 9 (August 16, 2018): 3153–69. http://dx.doi.org/10.1177/0954410018793288.

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A double shock waverider forebody configuration, with curved surfaces and known pressure fields and shock arrays, is constructed by a stream-tracing approach. The compression surface consists of a wedge and conical shocks. The conical shock results from a modified wave-derived bump surface that diverts the boundary layer before the inlet entrance. The design is fully computational fluid dynamics based and emphasis is placed on the compact design with boundary layer diverting ability. Controlling or diverting the thick boundary layer safely is a difficult challenge in hypersonic flight vehicle design especially when the inlets are highly integrated with the fuselage. Numerical simulations show that the new combination can divert a significant fraction of boundary layer before the inlet and maintains a good compression ratio for propulsion efficiency at Mach 5.0. Effects of forebody aerodynamics on the integrated inlet and comparisons with other systems are described in this paper.

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Cao, Yu Ji, Shi Ying Zhang, and Peng Gao. "Investigation of Attack Angle Character for Hypersonic Inlet." Advanced Materials Research 468-471 (February 2012): 1978–81. http://dx.doi.org/10.4028/www.scientific.net/amr.468-471.1978.

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The comparison analysis of performance with design point has been done in view of two different inlets, which were built by uniform shock intensity and shock angle method respectively. Base on this, the attack angle performance was emphatic developed on off-design point. The numerical calculation and analysis was conducted in eight different attack angle for air inlet conditions on two kinds of inlet respectively. The result indicated that, the inlet with uniform shock intensity method has 5% much more flow coefficient than inlet with uniform shock angle, in the condition of design point. In the condition of off-design point, the influence of the inlet performance is relatively small to the later kind of inlet with the increase of positive attack angel.

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Reddy, P. Nithish. "Hypersonic Scram-Jet Engine Inlet Design." International Journal for Research in Applied Science and Engineering Technology 7, no. 6 (June 30, 2019): 1619–35. http://dx.doi.org/10.22214/ijraset.2019.6273.

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Teng, Jian, and Hua Cheng Yuan. "Design Methodology and Unsteady Aerodynamic Characteristics of a Rectangular Variable Geometry Hypersonic Inlet." Applied Mechanics and Materials 275-277 (January 2013): 433–41. http://dx.doi.org/10.4028/www.scientific.net/amm.275-277.433.

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Design methodology of a rectangular variable geometry hypersonic inlet whose cowl lip is translatable along flow direction is clarified in current study and recommendation of key design parameters are given. Unsteady Reynolds-averaged Navier-Stokes (uRANS) calculation were carried out to investigate the feasibility and unsteady aerodynamic characteristics of this inlet. Results indicate that by stretching the movable lip of a model inlet upstream, mass flow rate will increases apparently due to the increases of inlet internal duct entrance area. Stretching the movable lip upstream will decrease CR of the model inlet which is favorable for the start or restart of the inlet from an unstarted status. The lip translating process is smooth and will not induce large amplitude flow disturbance within inlet duct. The movable lip is conducive to improve the aeropropulsive performance of the hypersonic inlet in wide flight range

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Li, Yuling, Lianjie Yue, Chengming He, Wannan Wu, and Hao Chen. "Lagrange Optimization of Shock Waves for Two-Dimensional Hypersonic Inlet with Geometric Constraints." Aerospace 9, no. 10 (October 20, 2022): 625. http://dx.doi.org/10.3390/aerospace9100625.

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The present paper focuses on the Lagrange optimization of shock waves for a two-dimensional hypersonic inlet by limiting the cowl internal angle and inlet length. The results indicate the significant influences of geometric constraints on the configuration of shock waves and performances of an inlet. Specifically, the cowl internal angle mainly affects the internal compression section; the inlet length affects both the internal and external compression sections where the intensity of internal and external compression shock waves shows a deviation of equal. In addition, the performances of optimized inlets at off-design points are further numerically simulated. A prominent discovery is that a longer inlet favors a higher total pressure recovery at the positive AOA; conversely, a shorter inlet can increase the total pressure recovery at the negative AOA.

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He, Xuzhao, Jialing Le, and Si Qin. "Design and analysis osculating general curved cone waverider." Aircraft Engineering and Aerospace Technology 89, no. 6 (October 2, 2017): 797–803. http://dx.doi.org/10.1108/aeat-12-2014-0214.

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Purpose Waverider has high lift to drag ratio and will be an idea aerodynamic configuration for hypersonic vehicles. But a structure permitting aerodynamic like waverider is still difficult to generate under airframe’s geometric constrains using traditional waverider design methods. And furthermore, traditional waverider’s aerodynamic compression ability cannot be easily adjusted to satisfy the inlet entrance requirements for hypersonic air-breathing vehicles. The purpose of this paper is to present a new method named osculating general curved cone (OCC) method aimed to improve the shortcomings of traditional waveriders. Design/methodology/approach A basic curved cone is, first, designed by the method of characteristics. Then the waverider’s inlet captured curve and front captured tube are defined in the waverider’s exit plane. Osculating planes are generated along the inlet captured curve and the designed curved cone is transformed to the osculating planes. Streamlines are traced in the transformed curved cone flow field. Combining all streamlines which have been obtained, OCC waverider’s compression surface is generated. Waverider’s upper surface uses the free stream surface. Findings It is found that OCC waverider has good volumetric characteristics and good flow compression abilities compared with the traditional osculating cone (OC) waverider. The volume of OCC waverider is 25 per cent larger than OC waverider at the same design condition. Furthermore, OCC waverider can compress incoming flow to required flow conditions with high total pressure recovery in the waverider’s exit plane. The flow uniformity in the waverider exit plane is quite well. Practical implications The analyzed results show that the OCC waverider can be a practical high performance airframe/forebody for hypersonic vehicles. Furthermore, this novel waverider design method can be used to design a structure permitting aerodynamic like waverider for a practical hypersonic vehicle. Originality/value The paper puts forward a novel waverider design method which can improve the waverider’s volumetric characteristics and compression abilities compared with the traditional waverider design methods. This novel design approach can extend the waverider’s applications for designing hypersonic vehicles.

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Chang, Juntao, Lei Wang, and Wen Bao. "Mathematical modeling and characteristic analysis of scramjet buzz." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 228, no. 13 (January 29, 2014): 2542–52. http://dx.doi.org/10.1177/0954410014521055.

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Buzz is an important issue for a scramjet engine. A mathematical model of buzz oscillations is necessary for control system design. Control-oriented models of hypersonic vehicle propulsion systems require a reduced-order model that is accurate to some extent but requires less than a few seconds of computational time. To achieve this goal, a reduced-order model of buzz oscillations for a scramjet engine is built by introducing the modeling idea of Moore–Greitzed model for compressors. The introduction of characteristic lines avoids the complex interactions in hypersonic inlet, such as shock–shock interactions and shock–boundary layer interaction. And the inlet characteristics are obtained from the pressure signal of combustor. Based on the established buzz model, we can predict the inlet performance, characterize the stability margin of inlet, reflect the oscillatory characteristics of inlet buzz including the dominant amplitude and frequency and describe the transition process of inlet buzz.

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Yan, Yilun, and Jiangfeng Wang. "Numerical Research on the NS-SDBD Control of a Hypersonic Inlet in Off-Design Mode." Aerospace 9, no. 12 (November 30, 2022): 773. http://dx.doi.org/10.3390/aerospace9120773.

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The overall performance of a scramjet inlet will decline while entering off-design mode. Active flow control using nanosecond surface dielectric barrier discharge (NS-SDBD) can be a novel solution to such inlet–unstart problems. NS-SDBD actuators are deployed on the surface of the internal compression section, controlling the shock waves and the separation area. Numerical simulations of hypersonic flows are carried out using the compressible Reynolds average Navier–Stokes equation (RANS), along with the plasma phenomenological model which is added in as the energy source term. Flow structures and the evolution of performance parameters are analyzed. Results show that NS-SDBD actuators are able to increase the static pressure behind the cowl shock, boosting the downstream total pressure. The compression effect becomes stronger while raising the frequency or shortening the spacing between the actuators. Under the inlet–unstart conditions, the compression wave generated by the actuator pushes the reattachment point forward, making the separation bubble longer in length and shorter in height, which reduces the strength of the separation shock. The results provide a numerical basis for the state control of a hypersonic inlet.

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10

Xiong, Bing, Xiao-qiang Fan, and Yi Wang. "Parameterization and optimization design of a hypersonic inward turning inlet." Acta Astronautica 164 (November 2019): 130–41. http://dx.doi.org/10.1016/j.actaastro.2019.07.004.

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Zhu, Chengxiang, Rijiong Yang, Rongqian Chen, Ruofan Qiu, and Yancheng You. "Investigation of adaptive slot control method for starting characteristics of hypersonic inlets." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 11 (December 25, 2018): 4261–71. http://dx.doi.org/10.1177/0954410018819523.

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Starting characteristics restrict the operation limits of a hypersonic inlet. Enhancement of the starting ability thus serves as one of the most serious issues in propulsion system. In the present work, we propose a simple adaptive slot control method, which expands the working range of hypersonic inlets to a lower Mach number and shows very weak losses. Our simulation results applying the five parallel slot geometrical design show a substantial reduction of the starting Mach number. The air flow inside the parallel slot channels is self-driven by the pressure gradient located near the separation shock under unstart mode, whereas it is strongly suppressed when the inlet is restarted. Surprisingly, all the inlet configurations are almost restarted at the same Mach 3.0, regardless of the individual width of the slot and the number of slot. This confirms the self-adapted nature of the pressure gradient inside the channel which shows prospect for the potential engineering applications of the simple slot control method. However, the location of the slots shows a big influence on the control efficiency, indicating that these slots need to be arranged carefully on the compression surface based on the location of the separation bubble.

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12

Yu, Zonghan, Guoping Huang, Ruilin Wang, and Omer Musa. "Spillage-Adaptive Fixed-Geometry Bump Inlet of Wide Speed Range." Aerospace 8, no. 11 (November 11, 2021): 340. http://dx.doi.org/10.3390/aerospace8110340.

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In this work, a new spillage-adaptive bump inlet concept is proposed to widen the speed range for hypersonic air-breathing flight vehicles. Various approaches to improve the inlet start-ability are summarized and compared, among which the bump-inlet pattern holds the merits of high lift-to-drag ratio, boundary layer diversion, and flexible integration ability. The proposed spillage-adaptive concept ensures the inlet starting performance by spilling extra mass flow away at low speed number conditions. The inlet presetting position is determined by synthetically evaluating the flow uniformity and the low-kinetic-energy fluid proportion. The numerical results show that the flow spillage of the inlet increases with the inflow speed decrease, which makes the inlet easier to start at low speed conditions (M 2.5–6.0). The effects of the boundary layer on spillage are also studied in this work. The new integration pattern releases the flow spillage potentials of three-dimensional inward-turning inlets by reasonably arranging the inlet compression on the bump surface. Future work will focus on the spillage-controllable design method.

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13

Xu, Shang-cheng, Yi Wang, Zhen-guo Wang, Xiao-qiang Fan, and Bing Xiong. "Design method for hypersonic bump inlet based on transverse pressure gradient." Journal of Zhejiang University-SCIENCE A 23, no. 6 (June 2022): 479–94. http://dx.doi.org/10.1631/jzus.a2100532.

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14

Kline, H. L., and J. J. Alonso. "Adjoint of Generalized Outflow-Based Functionals Applied to Hypersonic Inlet Design." AIAA Journal 55, no. 11 (November 2017): 3903–15. http://dx.doi.org/10.2514/1.j055863.

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15

You, YanCheng, and DeWang Liang. "Design concept of three-dimensional section controllable internal waverider hypersonic inlet." Science in China Series E: Technological Sciences 52, no. 7 (June 9, 2009): 2017–28. http://dx.doi.org/10.1007/s11431-009-0125-1.

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16

Wang, Jifei, Jinsheng Cai, Yanhui Duan, and Yuan Tian. "Design of shape morphing hypersonic inward-turning inlet using multistage optimization." Aerospace Science and Technology 66 (July 2017): 44–58. http://dx.doi.org/10.1016/j.ast.2017.02.018.

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Subbarao, Kamesh, and Jennifer D. Goss. "Combined Magnetohydrodynamic and Geometric Optimization of a Hypersonic Inlet." International Journal of Aerospace Engineering 2009 (2009): 1–12. http://dx.doi.org/10.1155/2009/793647.

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This paper considers the numerical optimization of a double ramp scramjet inlet using magnetohydrodynamic (MHD) effects together with inlet ramp angle changes. The parameter being optimized is the mass capture at the throat of the inlet, such that spillage effects for less than design Mach numbers are reduced. The control parameters for the optimization include the MHD effects in conjunction with ramp angle changes. To enhance the MHD effects different ionization scenarios depending upon the alignment of the magnetic field are considered. The flow solution is based on the Advection Upstream Splitting Method (AUSM) that accounts for the MHD source terms as well. A numerical Broyden-Flecher-Goldfarb-Shanno- (BFGS-) based procedure is utilized to optimize the inlet mass capture. Numerical validation results compared to published results in the literature as well as the outcome of the optimization procedure are summarized to illustrate the efficacy of the approach.

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Wang, Fang, and Shuang Lin Gao. "Numerical Study on Aerodynamic Design of Hypersonic Vehicle Forebody." Advanced Materials Research 756-759 (September 2013): 4626–29. http://dx.doi.org/10.4028/www.scientific.net/amr.756-759.4626.

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The lift forebody configuration of a small hypersonic vehicle is designed by using the wedge angle method in this paper. The lift forebody created has been optimized by the simplex method with a penalty function. The aerodynamic characteristics of the forebody optimized are investigated by numerical method. The research results show that the wedge angle method is a high efficient way to generate the lift forebody of the hypersonic vehicle; On the design mach number, there is pressure leaking between the upper and lower surface of lift forebody, which leads to lateral flow in the spanwise on the precompression plane, and which create the unhomogeneity of inlet flow field; Adding side skirts on the both sides, which can reduce the lateral flow on the forebody's precompression plane, it can raise the forebody lift.

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Zhu, Chengxiang, Xu Zhang, Fan Kong, and Yancheng You. "Design of a Three-Dimensional Hypersonic Inward-Turning Inlet with Tri-Ducts for Combined Cycle Engines." International Journal of Aerospace Engineering 2018 (November 27, 2018): 1–10. http://dx.doi.org/10.1155/2018/7459141.

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The operation of a propulsion system in terms of horizontal takeoff/landing and full-speed range serves as one of the main difficulties for hypersonic travelling. In the present work, a three-dimensional inward-turning inlet with tri-ducts for combined cycle engines is designed for the operation of three different modes controlled by a single rotational flap on the compression side, which efficiently simplifies the inlet structure and the flap control mechanism. At high flight speed between Mach 4 and 6, the pure scramjet mode is switched on, whereas both the ejector and the scramjet paths are open for a moderate Mach number between 2 and 4 with a larger throat area guaranteeing the inlet startability. In the low flight speed range with Mach number below 2, the additional turbojet path will be turned on to supply air for the turbine engine, whereas the other two paths remain open for spillage. Numerical simulations under different operation modes have proven the feasibility and good performance of the designed inlet, e.g., a nearly full mass flow ratio and a total pressure recovery around 0.5 can be achieved at the cruise speed. Meanwhile, the inlet works properly at low flight speeds which overcomes the typical starting problem of similar inlet designs. In the near future, wind tunnel experiments will be carried out to validate our inlet design and its performance.

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Ding, F., J. Liu, W. Huang, C. Peng, and S. Chen. "An airframe/inlet integrated full-waverider vehicle design using as upgraded aerodynamic method." Aeronautical Journal 123, no. 1266 (June 21, 2019): 1135–69. http://dx.doi.org/10.1017/aer.2019.49.

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ABSTRACTWith the aims of overcoming the limitations of the existing basic flow model derived from an axisymmetric generating body and extending the aerodynamic design method of the airframe/inlet integrated waverider vehicle, this study develops an upgraded basic flow model derived from an axisymmetric shock wave. It then upgrades the design method for airframe/inlet integration of an air-breathing hypersonic waverider vehicle, which is termed the ‘full-waverider vehicle’ in this study. In this paper, first, the design principle and method for the upgraded full-waverider vehicle derived from an axisymmetric basic shock wave are described in detail. Second, an upgraded basic flow model that accounts for both internal and external flows is derived from an axisymmetric basic shock wave by use of both the streamline tracing method and the method of characteristics (MOC). Third, the upgraded full-waverider vehicle is developed from the upgraded basic flow model by the streamline tracing method. Fourth, the design theories and methodologies of both the upgraded basic flow model and the upgraded full-waverider vehicle are validated by a numerical computation method. Finally, the aerodynamic performances and viscous effects of both the upgraded basic flow model and the upgraded full-waverider vehicle are analysed by numerical computation. The obtained results show that the upgraded basic flow model and aerodynamic design method are effective for the design of the airframe/inlet integration of an air-breathing hypersonic waverider vehicle.

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XU, Shangcheng, Yi WANG, Zhenguo WANG, Xiaoqiang FAN, and Xingyu ZHAO. "Design and analysis of a hypersonic inlet with an integrated bump/forebody." Chinese Journal of Aeronautics 32, no. 10 (October 2019): 2267–74. http://dx.doi.org/10.1016/j.cja.2019.04.010.

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De Vanna, Francesco, Danilo Bof, and Ernesto Benini. "Multi-Objective RANS Aerodynamic Optimization of a Hypersonic Intake Ramp at Mach 5." Energies 15, no. 8 (April 12, 2022): 2811. http://dx.doi.org/10.3390/en15082811.

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The work describes a systematic optimization strategy for designing hypersonic inlet intakes. A Reynolds-averaged Navier-Stokes database is mined using genetic algorithms to develop ideal designs for a priori defined targets. An intake geometry from the literature is adopted as a baseline. Thus, a steady-state numerical assessment is validated and the computational grid is tuned under nominal operating conditions. Following validation tasks, the model is used for multi-objective optimization. The latter aims at minimizing the drag coefficient while boosting the static and total pressure ratios, respectively. The Pareto optimal solutions are analyzed, emphasizing the flow patterns that result in the improvements. Although the approach is applied to a specific setup, the method is entirely general, offering a valuable flowchart for designing super/hypersonic inlets. Notably, because high-quality computational fluid dynamics strategies drive the innovation process, the latter accounts for the complex dynamics of such devices from the early design stages, including shock-wave/boundary-layer interactions and recirculating flow portions in the geometrical shaping.

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Sarosh, Ali, Dong Yun-Feng, and Muhammad Shoaib. "An Aerothermodynamic and Mass-Model Integrated Optimization Framework for Highly-Integrated Forebody-Inlet Configurations." Applied Mechanics and Materials 245 (December 2012): 277–82. http://dx.doi.org/10.4028/www.scientific.net/amm.245.277.

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A framework methodology for multidisciplinary multiobjective optimization and analysis is proposed. It is based on analytical aerothermodynamics and mass-modeling parameters of highly-integrated forebody-inlet configuration and representative hypersonic flight vehicle respectively. A complex configuration for a highly-integrated waverider forebody attached to planar compression ramps and planar sidewalled-inlet system is studied. Optimization and analytical solutions are obtained using SHWAMIDOF-FI design tool. Results show substantial improvement in geometric, performance and flow parameters as compared to baseline configuration.

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Miao, Hai Feng, Lv Rong Xie, and Weng Xiao Chai. "Effects of Unforced Bounday Layer Transition on the Performance of a Two-Dimensional Hypersonic Inlet." Applied Mechanics and Materials 275-277 (January 2013): 466–71. http://dx.doi.org/10.4028/www.scientific.net/amm.275-277.466.

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Numerical investigation was conducted on a typical two-dimensional hypersonic inlet to study the influence of unforced boundary layer transition affected by compression ramp geometric parameters on the inlet performance. The numerical results show that the transition onset location on the compression ramp can be delayed by filleting the ramp intersection, and also the inlet's performance obviously improves when the transition onset location is delayed. Compared with full turbulent situation, when the boundary layer transition occurs, the unstart of the inlet is significantly mitigated, the heat transfer rate on the compression ramp decreases, both the total pressure recovery coefficient and mass flow rate increase at both design and off-design points. But the static pressure distribution along the ramp is fairly independent of the varieties of boundary layer.

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Yu, Zonghan, Huihui Huang, Ruilin Wang, Yuedi Lei, Xueyang Yan, Zikang Jin, Omer Musa, and Guoping Huang. "Effects of Flow Spillage Strategies on the Aerodynamic Characteristics of Diverterless Hypersonic Inlets." Aerospace 9, no. 12 (November 29, 2022): 771. http://dx.doi.org/10.3390/aerospace9120771.

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This paper compares the aerodynamic characteristics of a central-spillage diverterless hypersonic inlet (i.e., bump inlet, Form 1) with a side-spillage inlet (Form 2) under on/off design conditions when faced with non-uniform inflow. Both forms are designed for a flight Mach number of 6.0 and a cruise altitude of 24.0 km. Numerical methods are introduced and validated. Integrated design results indicate that based on identical contraction ratios, Form 2 is 27.8% lower in height, 28.3% shorter in length, and 34.4% smaller in the windward projection area than Form 1. This provides the evidence that the side-spillage strategy will suppress the external drag less. Then, the aerodynamic performance is investigated under various upstream/downstream boundary conditions (inflow speed range: Mach 2.0~6.0; backpressure fluctuation range: 1~110.0 times the freestream static pressure). The evaluation methods for non-uniform flow fields are first introduced in this paper. Form 2 has a relatively stronger shock system, which allows it to suppress 4.52% more of the pressure fluctuation from the downstream combustion chamber than Form 1. The inlet start margin is widened by approximately 250% due to the self-adaptive flow spillage ability established by the side-spillage strategy. Furthermore, the compression efficiency, internal shock system, spillage ability, etc., are analyzed in detail. In summary, the side-spillage flow organization strategy has better potential for designing wide-ranging air-breathing flight vehicles.

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Kang, Sang Hun. "Variable Inlet Design for Hypersonic Engines with a Wide Range of Flight Mach Numbers." Journal of the Korean Society of Propulsion Engineers 19, no. 3 (June 1, 2015): 65–72. http://dx.doi.org/10.6108/kspe.2015.19.3.065.

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Wen, Xun, Jun Liu, Jie Li, Feng Ding, and Zhi-xun Xia. "Design and numerical simulation of a clamshell-shaped inlet cover for air-breathing hypersonic vehicles." Journal of Zhejiang University-SCIENCE A 20, no. 5 (May 2019): 347–57. http://dx.doi.org/10.1631/jzus.a1800620.

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Veeran, Sasha, Apostolos Pesyridis, and Lionel Ganippa. "Ramjet Compression System for a Hypersonic Air Transportation Vehicle Combined Cycle Engine." Energies 11, no. 10 (September 25, 2018): 2558. http://dx.doi.org/10.3390/en11102558.

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This report assesses the performance characteristics of a ramjet compression system in the application of a hypersonic vehicle. The vehicle is required to be self-powered and perform a complete flight profile using a combination of turbojet, ramjet and scramjet propulsion systems. The ramjet has been designed to operate between Mach 2.5 to Mach 5 conditions, allowing for start-up of the scramjet engine. Multiple designs, including varying ramp configurations and turbo-ramjet combinations, were investigated to evaluate their merits and limitations. Challenges arose with attempting to maintain sufficient pressure recoveries and favourable flow characteristics into the ramjet combustor. The results provide an engine inlet design capable of propelling the vehicle between the turbojet and scramjet phase of flight, allowing for the completion of its mission profile. Compromises in the design, however, had to be made in order to allow for optimisation of other propulsion systems including the scramjet nozzle and aerodynamics of the vehicle; it was concluded that these compromises were justified as the vehicle uses the ramjet engine for a minority of the flight profile as it transitions between low supersonic to hypersonic conditions.

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Ji, Zifei, Huiqiang Zhang, and Bing Wang. "Thrust control strategy based on the minimum combustor inlet Mach number to enhance the overall performance of a scramjet engine." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 13 (February 20, 2019): 4810–24. http://dx.doi.org/10.1177/0954410019830816.

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A lower combustor inlet Mach number is desirable in order to design a compact, lightweight combustor and boost the overall performance of the scramjet engine. In this study, a thrust control strategy is proposed for a hydrogen-fueled scramjet taking into account the operating limitations, which is called the minimum combustor inlet Mach number rule since the combustor inlet Mach number is used as the control variable. By scheduling the fuel supply and modifying the intake geometry, the combustor inlet Mach number can be minimized while ensuring a certain thrust output within the operation constraints. In this manner, the scramjet engine can be operated with high specific thrust and low fuel consumption throughout the flight envelope. The thrust control strategy is further applied to a hydrogen-fueled scramjet in the hypersonic flight regime. Because the combustor inlet Mach number varies with flight conditions, the thrust strategy can be applied in practice by monitoring the following aerothermodynamic parameters in different flight regimes instead: (1) combustor outlet Mach number, (2) combustor inlet static temperature, and (3) combustor outlet static temperature. Furthermore, the effects of the thrust output on the division of flight regime are investigated, and the overall performance of the hydrogen-fueled scramjet engine obtained from applying the thrust control strategy is discussed in detail.

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Utomo, Muhammad Adnan, and Romie Oktavianus Bura. "Design of Inward-Turning External Compression Supersonic Inlet for Supersonic Transport Aircraft." INSIST 2, no. 2 (January 25, 2019): 104. http://dx.doi.org/10.23960/ins.v2i2.90.

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Inward-turning external compression intake is one of the hybrid intakes that employs both external and internal compression intakes principle. This intake is commonly developed for hypersonic flight due to its efficiency and utilizing fewer shockwaves that generate heat. Since this intake employ less shockwaves, this design can be applied for low supersonic (Mach 1.4 - 2.5) intakes to reduce noise generated from the shockwaves while maintaining the efficiency. Other than developing the design method, a tool is written in MATLAB language to generate the intake shape automatically based on the desired design requirement. To investigate the intake design tool code and the performance of the generated intake shape, some CFD simulation were performed. The intake design tool code can be validated by comparing the shockwave location and the air properties in every intake's stations. The performance parameters that being observed are the intake efficiency, flow distortion level at the engine face, and the noise level generated by the shockwaves. The design tool written in MATLAB is working as intended. Two dimensional axisymmetric CFD simulations validation has been done and the design meets the minimum requirement. As for the 3D inlet geometry, with a little modification on diffuser and equipping vent to release the buildup pressure, the inlet has been successfully met the military standard on inlet performance (MIL-E-5007D). This design method also has feature to fit every possible throat cross sectional shapes and has been proven to work as designed.Keywords— Inward-turning, Supersonic, Engine Intakes, Low- noise, Design Method

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Ding, Feng, Jun Liu, Chi-bing Shen, Wei Huang, Zhen Liu, and Shao-hua Chen. "An overview of waverider design concept in airframe/inlet integration methodology for air-breathing hypersonic vehicles." Acta Astronautica 152 (November 2018): 639–56. http://dx.doi.org/10.1016/j.actaastro.2018.09.002.

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Sarosh, Ali, Dong Yun Feng, and Muhammad Adnan. "An Aerothermodynamic Design Approach for Scramjet Combustors and Comparative Performance of Low-Efficiency Systems." Applied Mechanics and Materials 110-116 (October 2011): 4652–60. http://dx.doi.org/10.4028/www.scientific.net/amm.110-116.4652.

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This paper is aimed at development of an integrated approach based on analytical and computational aerothermodynamics for the special case of design of a 75% (low process-efficiency), hydrogen-fuelled, constant area combustor of a hypersonic airbreathing propulsion (HAP) system thereafter undertaking study of two types of HAP systems. The results of configurational aerothermodynamics implied that the most appropriate constant area configuration had a 30 degrees downstream wall-mounted fuel injector with a single acoustically stable cavity placed downstream of the fuel injection point. Moreover for identical flow inlet parameters and system configurations at lower levels of thermodynamic process efficiencies, the constant combustor-area (i.e. Scramjet 1) engine is superior in its performance to the constant combustor-pressure (i.e. Scramjet 2) engine for all values of fuel-air ratios.

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Wang, Chengpeng, Xin Yang, Longsheng Xue, Konstantinos Kontis, and Yun Jiao. "Correlation Analysis of Separation Shock Oscillation and Wall Pressure Fluctuation in Unstarted Hypersonic Inlet Flow." Aerospace 6, no. 1 (January 10, 2019): 8. http://dx.doi.org/10.3390/aerospace6010008.

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The flow field in a hypersonic inlet model at a design point of M = 6 has been studied experimentally. The focus of the current study is to present the time-resolved flow characteristics of separation shock around the cowl and the correlation between the separation shock oscillation induced by the unstart flow and the wall pressure fluctuation when the inlet is in a state of unstart. High-speed Schlieren flow visualization is used to capture the transient shock structure. High-frequency pressure transducers are installed on the wall around both the cowl and isolator areas to detect the dynamic pressure distribution. A schlieren image quantization method based on gray level detection and calculation is developed to analyze the time-resolved spatial structure of separation shock. Results indicate that the induced separation shock oscillation and the wall pressure fluctuation are closely connected, and they show the same frequency variation characteristics. The unsteady flow pattern of the “little buzz” and “big buzz” modes are clarified based on time-resolved Schlieren images of separation shock. Furthermore, the appropriate location of the pressure transducers is determined on the basis of the combined analysis of fluctuating wall-pressure and oscillating separation shock data.

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34

Wood, J. R., J. F. Schmidt, R. J. Steinke, R. V. Chima, and W. G. Kunik. "Application of Advanced Computational Codes in the Design of an Experiment for a Supersonic Throughflow Fan Rotor." Journal of Turbomachinery 110, no. 2 (April 1, 1988): 270–79. http://dx.doi.org/10.1115/1.3262191.

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Increased emphasis on sustained supersonic or hypersonic cruise has revived interest in the supersonic throughflow fan as a possible component in advanced propulsion systems. Use of a fan that can operate with a supersonic inlet axial Mach number is attractive from the standpoint of reducing the inlet losses incurred in diffusing the flow from a supersonic flight Mach number to a subsonic one at the fan face. The data base for components of this type is practically nonexistent; therefore, in order to furnish the required information for assessment of this type fan, a program has been initiated at the NASA Lewis Research Center to design, build, and test a fan rotor that operates with supersonic axial velocities from inlet to exit. This paper describes the design of the experiment using advanced computational codes to calculate the unique components required. The fan rotor has constant hub and tip radii and was designed for a pressure ratio of 2.7 with a tip speed of 457 m/s. The rotor was designed using existing turbomachinery design and analysis codes modified to handle fully supersonic axial flow through the rotor. A two-dimensional axisymmetric throughflow design code plus a blade element code were used to generate fan rotor velocity diagrams and blade shapes. A quasi-three-dimensional, thin shear layer Navier–Stokes code was used to assess the performance of the fan rotor blade shapes. The final design was stacked and checked for three-dimensional effects using a three-dimensional Euler code interactively coupled with a two-dimensional boundary layer code. A translating nozzle was designed to produce a uniform flow parallel to the fan up to the design axial Mach number of 2.0. The nozzle was designed with the three-dimensional Euler/interactive boundary layer code. The nozzle design in the expansion region was analyzed with a three-dimensional parabolized viscous code, which corroborated the results from the Euler code. A translating supersonic diffuser was designed using these same codes.

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35

Navó, Àlex, and Josep M. Bergada. "Aerodynamic Study of the NASA’s X-43A Hypersonic Aircraft." Applied Sciences 10, no. 22 (November 19, 2020): 8211. http://dx.doi.org/10.3390/app10228211.

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A 2D aerodynamic study of the NASA’s X-43A hypersonic aircraft is developed using two different approaches. The first one is analytical and based on the resolution of the oblique shock wave and Prandtl–Meyer expansion wave theories supported by an in-house program and considering a simplified aircraft’s design. The second approach involves the use of a Computational Fluid Dynamics (CFD) package, OpenFOAM and the real shape of the aircraft. The aerodynamic characteristics defined as the lift and drag coefficients, the aerodynamic efficiency and the pitching moment coefficient are calculated for different angles of attack. Evaluations are made for an incident Mach number of 7 and an altitude of 30 km. For both methodologies, the required angles of attack to achieve a Vertical Force Balance (VFB) and a completely zero pitching moment conditions are considered. In addition, an analysis to optimise the nose configuration of the aircraft is performed. The mass flow rate throughout the scramjet as a function of the angle of attack is also presented in the CFD model in addition to the pressure, density, temperature and Mach fields. Before presenting the corresponding results, a comparison between the aerodynamic coefficients in terms of the angle of attack of both models is carried out in order to properly validate the CFD model. The paper clarifies the requirements needed to make sure that both oblique shock waves originating from the leading edge meet just at the scramjet inlet clarifying the advantages of fulfilling such condition.

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Al-Garni, A. Z., A. Z. Şahin, and B. S. Yilbas. "Active Cooling of a Hypersonic Plane Using Hydrogen, Methane, Oxygen and Fluorine." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 210, no. 1 (January 1996): 9–17. http://dx.doi.org/10.1243/pime_proc_1996_210_340_02.

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This paper studies active cooling of an aerospace plane using liquid hydrogen, liquid methane, liquid oxygen and liquid fluorine. An ascending optimized trajectory to minimize the heat load in the hypersonic part is used to perform the study, which includes cooling of the stagnation point, the leading edges of wings and engine and other parts of the aerospace plane that are close to the leading edges. The laminar case of the stagnation point and both laminar and turbulent cases for the leading edge heating have been considered. The amount of liquid coolant mass needed for cooling is calculated. A design of minimum inlet–outlet areas for the amount of liquid needed for cooling is made with consideration of the coolant's physical constraints in the liquid and gaseous states. The study shows that the ratio of masses of coolant to the initial total mass (initial total mass of the vehicle including fuel and coolant masses) is in the limit of the reachable range. The comparison shows that the hydrogen is a clear winner as a candidate for coolant and saves mass as compared to the other three coolants. The study shows that there are no fundamental barriers for the cooling system of the vehicle in terms of its coolant mass and area size for coolant passage.

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Fan, Y., J. Chang, W. Bao, and D. Yu. "Effects of boundary-layer bleeding on unstart oscillatory flow of hypersonic inlets." Aeronautical Journal 114, no. 1157 (July 2010): 445–50. http://dx.doi.org/10.1017/s0001924000003924.

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Abstract The unsteady flowfield of a series of mixed-compression hypersonic inlets with different bleeding rates were numerically simulated. Firstly unstart oscillatory flow of hypersonic inlets caused by downstream massflow choking was discussed. Then the effects of boundary layer bleeding on the averaged performance parameter of hypersonic inlets, and on the dominant amplitude and frequency of unstart oscillatory flow of hypersonic inlets were presented. The reasons why the boundary-layer bleeding can suppress unstart oscillatory flow of hypersonic inlets were analysed. In conclusion, the averaged performance parameter of hypersonic inlets during a big buzz is improved greatly, and the dominant frequency of unstart oscillatory flow of hypersonic inlets is reduced in contrast with no bleeding, and all these are benefit to the design and operation of hypersonic inlets.

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38

Lin, Sheam-Chyun, and Yu-Shan Luo. "Integrated design of hypersonic waveriders including inlets and tailfins." Journal of Spacecraft and Rockets 32, no. 1 (January 1995): 48–54. http://dx.doi.org/10.2514/3.26573.

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39

Shukla, Vijay, Andrew Gelsey, Mark Schwabacher, Donald Smith, and Doyle D. Knight. "Automated Design Optimization for the P2 and P8 Hypersonic Inlets." Journal of Aircraft 34, no. 2 (March 1997): 228–35. http://dx.doi.org/10.2514/2.2161.

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Raj, N. Om Prakash, and K. Venkatasubbaiah. "A new approach for the design of hypersonic scramjet inlets." Physics of Fluids 24, no. 8 (August 2012): 086103. http://dx.doi.org/10.1063/1.4748130.

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41

Gollan, Rowan J., and Michael K. Smart. "Design of Modular Shape-Transition Inlets for a Conical Hypersonic Vehicle." Journal of Propulsion and Power 29, no. 4 (July 2013): 832–38. http://dx.doi.org/10.2514/1.b34672.

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42

Wang, Chengpeng, Xuang Tian, Lingfeng Yan, Longsheng Xue, and Keming Cheng. "Preliminary Integrated Design of Hypersonic Vehicle Configurations Including Inward-Turning Inlets." Journal of Aerospace Engineering 28, no. 6 (November 2015): 04014143. http://dx.doi.org/10.1061/(asce)as.1943-5525.0000480.

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43

Lu, H., and D. Cao. "Three-dimensional design and numerical simulation of magnetohydrodynamic controlled hypersonic inlets." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 226, no. 8 (October 13, 2011): 1002–13. http://dx.doi.org/10.1177/0954410011416059.

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Smart, M. K. "Design of Three-Dimensional Hypersonic Inlets with Rectangular-to-Elliptical Shape Transition." Journal of Propulsion and Power 15, no. 3 (May 1999): 408–16. http://dx.doi.org/10.2514/2.5459.

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45

WANG, Jifei, Jinsheng CAI, Chuanzhen LIU, Yanhui DUAN, and Yaojie YU. "Aerodynamic configuration integration design of hypersonic cruise aircraft with inward-turning inlets." Chinese Journal of Aeronautics 30, no. 4 (August 2017): 1349–62. http://dx.doi.org/10.1016/j.cja.2017.05.002.

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46

Cui, Kai, ShouChao Hu, GuangLi Li, ZhiPeng Qu, and Ming Situ. "Conceptual design and aerodynamic evaluation of hypersonic airplane with double flanking air inlets." Science China Technological Sciences 56, no. 8 (June 28, 2013): 1980–88. http://dx.doi.org/10.1007/s11431-013-5288-0.

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47

Fu, Lei, Shuai Zhang, and Yao Zheng. "Performances analysis of asymmetric minimum length nozzles." International Journal of Modeling, Simulation, and Scientific Computing 07, no. 02 (June 2016): 1650021. http://dx.doi.org/10.1142/s1793962316500215.

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Two-dimensional (2D) and axisymmetric minimum length nozzles (MLNs) with constant and variable specific heat were strictly designed using the method of characteristics (MOCs). MOC is a numerical technique which has great advantages in accuracy and efficiency for solving hyperbolic partial differential equations. According to previous MLN designs, violent vibrations of upper wall discrete points at the inlet were observed for 2D nozzles. Meanwhile, slight compressions could be observed in the flow field of axisymmetric nozzles designed by those methods. We proposed a novel technique in which the inlet grid is intensified to overcome the limitations mentioned above. Inviscid numerical simulations by CFD revealed that the proposed nozzle could meet the requirements for exit Mach number and flow field uniformity. Additionally, asymmetric MLNs could be used to hypersonic vehicles. The preliminary performances of 2D asymmetric nozzle with constant specific heat were investigated.

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48

Kovalchuk, O., O. Skorohvatov, A. Galkin, L. Gordishevski, and V. Liskovchuk. "ARMORED SHELLS WITH STRAIGHT AIR JET ENGINE." Collection of scientific works of Odesa Military Academy 1, no. 13 (December 30, 2020): 34–43. http://dx.doi.org/10.37129/2313-7509.2020.13.1.34-43.

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The article analyzes the problems of sub-caliber feathered projectile, offers a variant of equipping such a projectile with an air-jet engine, graphs of air resistance, calculations of the required thrust of the air-jet engine. Features of armor-piercing projectiles with a direct-flow air-jet engine are considered, the main calculations are performed using high-level Python programming language. Currently, as armor-piercing ammunition are widely used armor-piercing sub-caliber feathered shells (BPOS) with high penetrating ability This is achieved due to the high initial velocity of ammunition (1650-1840 m / s) and small cross section (d = 20-30 mm). To compensate for the force of air resistance, the provision of jet propulsion ammunition is used. But the main disadvantage of such shells is the dependence of the ability to pierce armor from a distance to the target. That is, due to the resistance of the air, the speed of the projectile is lost, namely its energy. What they are inferior to cumulative projectiles, for which the ability to pierce armor does not depend on the distance to the target. Modern armored vehicles have significant armor and BPOS lose their importance in the range of cumulative projectiles and anti-RPG. This situation can be corrected if the BPOS is equipped with direct-flow jet engines (PPD). Direct-flow air jet engine (PPD), simple in design, has a high efficiency at large Mach numbers, compact, because it does not require the presence of an oxidant in the fuel, as it uses oxygen from the environment. Compressed air entering the combustion chamber from the inlet device is heated by oxidation of the fuel supplied to it. Created from a mixture of air with combustion products gas mixture – the working fluid in the nozzle reaches the speed of sound, and at its output expanding to supersonic. The working fluid flows at a speed greater than the speed of the oncoming air flow, which creates a jet thrust. When the flight speed is much less than the speed of the jet, the thrust increases. As the speed of flight approaches the speed of the jet, the thrust decreases, passing some maximum corresponding to the optimal speed of flight. With the development of mixed solid fuel technology, it began to be used in PPRD. A fuel checker with a longitudinal Central channel is placed in the combustion chamber. The working fluid passing through the combustion chamber oxidizes the fuel from its surface and heats up. The use of solid fuel further simplifies the design of the PPRD as it does not require a combustion chamber. The main part of the filler of mixed fuel PPRD is a fine powder of aluminum, magnesium or beryllium, the heat of combustion, which is much higher than the heat of combustion of hydrocarbon fuels. With the development of mixed solid fuel technology, it began to be used in PPRD. A fuel checker with a longitudinal Central channel is placed in the combustion chamber. The working fluid passing through the combustion chamber oxidizes the fuel from its surface and heats up. The use of solid fuel further simplifies the design of the PPRD as it does not require a combustion chamber. The main part of the filler of mixed fuel PPRD is a fine powder of aluminum, magnesium or beryllium, the heat of combustion, which is much higher than the heat of combustion of hydrocarbon fuels. An example of a solid propellant PPRD can be the propulsion engine of the anti-ship missile P-270 Mosquito. Depending on the speed of flight PPRD are divided into subsonic, supersonic and hypersonic. This division is due to the design features of each of these groups. In the supersonic range PPRD is much more effective than in the subsonic. For example, at a speed of M = 3, the degree of pressure increase in the PPRD is 37, which can be compared with the most high-pressure compressors of turbojet engines. Keywords: armor-piercing sub-caliber feathered projectile, air-jet engine, external ballistics.

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Brahmachary, Shuvayan, Ganesh Natarajan, Vinayak Kulkarni, and Niranjan Sahoo. "Comment on “A new approach for the design of hypersonic scramjet inlets” [Phys. Fluids 24, 086103 (2012)]." Physics of Fluids 32, no. 7 (July 1, 2020): 079101. http://dx.doi.org/10.1063/5.0006408.

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Prakash Raj, N. Om, and K. Venkatasubbaiah. "Response to “Comment on ‘A new approach for the design of hypersonic scramjet inlets’” [Phys. Fluids 32, 079101 (2020)]." Physics of Fluids 32, no. 7 (July 1, 2020): 079102. http://dx.doi.org/10.1063/5.0012513.

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Journal articles on the topic 'Hypersonic inlet design'

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