Mach 5 inlet CFD and experimental results

Weir, Lois J.; Reddy, D. R.; Rupp, George D.

doi:10.2514/6.1989-2355

Cited by 20 publications

(4 citation statements)

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“…The maximum recovery factor reaches to 0.5 and 0.56 for the cases with and without throat bleed. These values are considerably high comparing to the other results [7] and almost as high as the MIL-E-5008B target. Since the entry Mach number was 4.5, that is lower than Mach 5, the theoretical shock loss at designed 1st ramp, (5.1deg), which is 0.978, should be multiplied to the results for comparing the model performance to MIL-E-5008B.…”

Section: Resultsmentioning

confidence: 50%

Experimental Study of Mixed Compression Air-Intake for Hypersonic Airbreathing Engines

Sakata

Yanagi

Murakami

et al. 1992

Volume 2: Aircraft Engine; Marine; Microturbines and Small Turbomachinery

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Supersonic air-intake for Mach number higher than 2.5 is being investigated with experimental, analytical and computational methods. The study is performed in a part of the joint research program led by National Aerospace Laboratory (NAL) on the hypersonic airbreathing turbo-engines with subsonic ram combustion. The wind tunnel models are designed in two-dimensional mixed compression type with multi-shock system and tested in NAL’s Mach 4 supersonic wind tunnel. Pressure measurements and flow visualization by schlieren method, oil-flow and vapor screen techniques are being done. Here, the test results of Mach 4 and Mach 5 models are discussed. The Mach 4 model is fixed geometry with 5-shock system and the Mach 5 one is variable geometry with 6-shock and an isentropic compression surface. An expansion fore-plate was installed at the Mach 5 model inlet to accelerate the air-speed at the entry. The bleed systems at throat, ramp and cowl are adopted and evaluated in terms of pressure recovery and stability. Importance of establishment of the internal shock wave system, reduction of upstream Mach number of terminal shock wave and suppression of flow separation at diffusers are found. It is also found that ramp bleed is effective to confirm intake start and to minimize shock/boundary layer interaction.

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Section: Resultsmentioning

confidence: 50%

Experimental Study of Mixed Compression Air-Intake for Hypersonic Airbreathing Engines

Sakata

Yanagi

Murakami

et al. 1992

Volume 2: Aircraft Engine; Marine; Microturbines and Small Turbomachinery

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“…Based on the oblique shock solution to the Rankine-Hugoniot equations, we obtain the cowl compression angle α 1 and the Mach number Ma cowl before the cowl-induced incident shock. Statistically speaking, α 1 is usually 10 • -15 • (Weir et al 1989;Rodi, Emami & Trexler 1996;Schmitz & Bissinger 1998;Sanders & Weir 1999;Taguchi et al 2003;Albertson, Emami & Trexler 2006;Sanders & Weir 2008;Chang et al 2012;Tan et al 2012;Gounko, Mazhul & Nurutdinov 2014;Zhang et al 2015Zhang et al , 2016You et al 2017;Zhang et al 2017;Devaraj et al 2020;Huang et al 2021;Saravanan et al 2021), as shown in figure 2.…”

Section: Description Of Simplified Inlet Modelmentioning

confidence: 99%

“…However, to achieve enhanced flow compression for drag reduction purposes, the flow turning angle at the inlet internal cowl is typically maintained above 10 • (Weir, Reddy & RUPP 1989;Devaraj et al 2020;Huang et al 2021). In recent years, researchers have shifted their focus towards more practical investigations.…”

Section: Introductionmentioning

confidence: 99%

On incident shock wave/boundary layer interactions under the influence of expansion corner

Guo,

Tan,

Zhang

et al. 2024

J. Fluid Mech.

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Cowl-induced incident shock wave/boundary layer interactions (ISWBLIs) under the influence of shoulder expansion represent one of the dominant phenomena in supersonic inlets. To provide a more comprehensive understanding of how an expansion corner affects the ISWBLI, a detailed experimental and analytical study is performed in a Mach 2.73 flow in this work. Pressure measurement, schlieren photography and surface oil-flow visualisation are used to record flow features, including the pressure distribution, separation extent and surface-flow topological structures. Our results reveal three types of ISWBLIs influenced by the expansion corner. When the shock intensity is weak, the separation is small scale with the expansion waves emanating from the expansion corner. This is the first type of expansion-corner-affected ISWBLI (EC-ISWBLI). When the incident shock wave is strong, large-scale separation occurs, accompanied by the disappearance of expansion waves, forming the second type of EC-SWBLI. The expansion corner induces a ‘lock-in’ effect in which the separation onset is consistently locked near the expansion corner regardless of the incident shock intensity and impingement position. The third type of EC-ISWBLI occurs when the shock is sufficiently strong and the impingement point is close to the expansion corner. In this interaction, the ‘lock-in’ effect ceases to manifest. Moreover, a shock polar-incorporating inviscid model is employed to elucidate the shock patterns. Two criteria are established by combining free interaction theory with this model. The first criterion provides valuable insights into the evolution of separations with a minimal overall pressure rise and the second criterion determines the threshold for the occurrence of the ‘lock-in’ effect.

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“…In general, hypersonic vehicles operate at very high altitudes yielding a higher service ceiling encounter with very low air pressure and density (Fan 2011). The required pressure rise at a very high altitude is achieved by mixed compression inlets (Van Wie et al 1996) where the required compression is achieved by the external compression from the compression ramp and through a system of multiple shocks inside the 634 duct (Weir et al 1989). In a supersonic inlet, there are various flow phenomena such as shock waves and their interaction, internal and external shocks, and boundary layer separation.…”

Section: Introductionmentioning

confidence: 99%

Flow Characteristics of a Mixed Compression Hypersonic Intake

2022

JAFM

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The flow field in a two-dimensional hypersonic mixed-compression inlet in a freestream Mach numbers of M∞ =2.0, 3.0, and 5.0 are numerically solved to understand the effect of throat area variation. The exit area ratio variation is simulated by placing a plug insert at different axial locations at the exit of the model. The flow field is achieved computationally by solving the Reynolds Averaged Navier-Stokes equations in a finite volume framework. For each flow condition, the variation in shock structure is analyzed and the variation of the oblique shock wave angle with the mass flow rate is calculated theoretically and compared with the present CFD analysis. The variation in oblique shock angle is calculated in terms of the mass flow rate by considering the capture area and spillage flow through the inlet. The theoretical results suggest that the method can predict the inlet operating conditions at different freestream Mach numbers and area ratios. This method can quantify the reduction in mass flow rate due to the throttling effect by analyzing the flow field shock pattern. The effects of various important performance parameters such as free stream Mach number, total pressure recovery, and mass flow ratio were then numerically investigated. As the Mach number is increased, the total pressure recovery is reduced, but the maximum value of the mass flow rate is increased. The analysis is also focused on the effect of throat area variation on performance parameters at each Mach number. The characteristic curve of the inlet is then obtained for each free stream Mach number.

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Mach 5 inlet CFD and experimental results

Cited by 20 publications

References 0 publications

Experimental Study of Mixed Compression Air-Intake for Hypersonic Airbreathing Engines

Experimental Study of Mixed Compression Air-Intake for Hypersonic Airbreathing Engines

On incident shock wave/boundary layer interactions under the influence of expansion corner

Flow Characteristics of a Mixed Compression Hypersonic Intake

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