Numerical and experimental study on waverider with blunt leading edge

Liu, Jianxia; Hou, Zhanqiang; Ding, Guanghong; Chen, Xiaoqing; Chen, Xiaoqian

doi:10.1016/j.compfluid.2013.06.005

Cited by 23 publications

(5 citation statements)

References 32 publications

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“…26 Details of this tunnel have been reported by Liu et al. 24,28 The boundary conditions for the free stream are set as a Mach number of 8.04, a Reynolds number per unit of 1.13 × 10 7 /m, a total pressure of 7.8 MPa, and a total temperature of 892 K, in accordance to the experimental setup in Li. 26 The detailed aerodynamic performance data of the space shuttle measured during the wind tunnel experiment 26 is listed in Table 4.…”

Section: Physical Model and Numerical Methodsmentioning

confidence: 99%

Comparison between novel waverider generated from flow past a pointed von Karman ogive and conventional cone-derived waverider

Ding

Shen

Liu

et al. 2015

Proceedings of the Institution of Mechanical Engineers, Part G:

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This study conducts a comparison between a novel waverider generated from axisymmetric supersonic flow past a pointed von Karman ogive and a conventional cone-derived waverider generated from the axisymmetric supersonic flow past a cone. First, a pointed von Karman ogive is obtained by sharpening the blunt nose of the original von Karman ogive to a point. Then, a cone with the same shock wave radius at the base plane as that of the pointed von Karman ogive when it travels at a designated supersonic speed is designed. Both axisymmetric supersonic flows past the pointed von Karman ogive and the cone are calculated using the method of characteristics, and the two calculated flow fields are employed as the new and conventional basic flow fields of the waveriders, respectively. Second, a novel waverider is generated from the new basic flow field past the pointed von Karman ogive, using the streamline tracing technique, and a conventional cone-derived waverider is generated from the conventional basic flow field past the cone, using the streamline tracing technique. Finally, numerical methods are employed to verify the design procedures of the novel and conventional cone-derived waveriders, and to analyze the differences between both types of waverider. The obtained results show that the novel waverider possesses lower drag and higher lift-to-drag ratios, while the conventional cone-derived waverider has a higher internal volume and volumetric efficiency.

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Section: Physical Model and Numerical Methodsmentioning

confidence: 99%

Comparison between novel waverider generated from flow past a pointed von Karman ogive and conventional cone-derived waverider

Ding

Shen

Liu

et al. 2015

Proceedings of the Institution of Mechanical Engineers, Part G:

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show abstract

“…Sharp leading edges are becoming an increasingly desirable feature on hypersonic vehicles to improve their aerodynamic characteristics for both re-entry [1] and sustained hypersonic flight [2,3]. The improvements in lift-to-drag ratio of vehicles with sharp leading edges have been explored for hypersonic vehicles [4][5][6], where appreciable aerodynamic benefits are found for sharper cases. In these studies, cylindrically blunted leading edges with 𝑅 𝐿𝐸 ranging from 0.1 mm to 30 mm were explored.…”

Section: Introductionmentioning

confidence: 99%

Aerodynamic Effects and Heat Flux Augmentation of a Transpiration Cooled Hypersonic Sharp Leading Edge

Ravichandran

Doherty

McGilvray

et al. 2023

AIAA SCITECH 2023 Forum

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This paper presents numerical results investigating the aerodynamic and aerothermal effects of mass injection applied to hypersonic sharp leading edges, in the context of active thermal protection systems. A numerical study was carried out using Eilmer to investigate the coupling of leading edge radius and mass injection on heat flux and drag augmentations. Radii from 1 mm to 25 mm were considered at blowing parameters from 0.0 to 1.5 on 2D planar leading edge at a fixed trajectory point. Drag was found to barely change with mass injection, whereas heat flux was found to significantly reduce at the leading edge. The leading edge heat flux distribution could be collapsed, and therefore predicted, straightforwardly. Film cooling predictive methods in literature for other mass injection scenarios were found to collapse the heat flux and concentration comfortably with empirical modifications.

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“…Under hypersonic flight conditions, a large temperature gradient is formed along the direction of the airflow, and a large thermal stress is generated in the high-temperature area of the front end of the sharp wedge structure, which makes the front end of the wedge structure face a serious problem of excessive stress, and fullscale thermal testing of the structure is needed [10][11][12][13][14][15][16][17][18]. However, the traditional quartz lamp and graphite heating equipment are limited by their heating materials, and there are shortcomings in simulating temperature or continuous working time [19][20][21][22][23][24]. In addition, the operating cost of arc wind tunnels is very high, so gas flow wind tunnel testing is a good choice for conducting thermal tests [25][26][27][28][29].…”

Section: Introductionmentioning

confidence: 99%

Study on Ground Experimental Method of Stagnation Point Large Heat Flux of Typical Sharp Wedge Leading Edge Structure

Wang,

Lou,

et al. 2023

Aerospace

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In this paper, aimed at the problem of large temperature gradient thermal testing with the typical sharp wedge leading edge structure of a hypersonic vehicle, a subsonic high-temperature combustion gas heating (SHCH) test device is used to conduct a series of experiments on the heat flux simulation ability of subsonic high-temperature combustion gas in the stagnation point region. Firstly, for a hypersonic vehicle with a flying height of 24 km and Mach number range of 4~6.5, the stagnation point heat flux in the head area is obtained by numerical calculation of a typical leading edge structure, which is used as the experimental target of the thermal structure test. Secondly, an experimental specimen with a Gardon heat flux meter is designed with the same shape and size as the specimen in the numerical simulations to prepare for the subsequent SHCH test. Thirdly, a method to determine the combustion gas temperature based on a Kriging surrogate model is proposed. CFD numerical simulation is conducted using the SHCH test model, and the numerical calculation results are used as the training dataset. The Kriging surrogate model is used to establish an approximate fitting relationship between the stagnation point heat flux and experimental parameters under SHCH conditions. The corresponding combustion gas temperature values are found, respectively, with the hypersonic aerodynamic heat flux at Mach 5.0~5.4 as the target value. Finally, stagnation point heat flux testing of low-speed and high-temperature combustion gas is performed at different combustion gas temperatures. The experimental and target values obtained from hypersonic aerodynamic thermal simulations are compared and analyzed to verify the heating capacity of SHCH and the feasibility of hypersonic aerothermal simulation testing.

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Numerical and experimental study on waverider with blunt leading edge

Cited by 23 publications

References 32 publications

Comparison between novel waverider generated from flow past a pointed von Karman ogive and conventional cone-derived waverider

Comparison between novel waverider generated from flow past a pointed von Karman ogive and conventional cone-derived waverider

Aerodynamic Effects and Heat Flux Augmentation of a Transpiration Cooled Hypersonic Sharp Leading Edge

Study on Ground Experimental Method of Stagnation Point Large Heat Flux of Typical Sharp Wedge Leading Edge Structure

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