scholarly journals A low-boom and low-drag design method for supersonic aircraft and its applications on airfoils

2021 ◽  
Vol 3 (1) ◽  
Author(s):  
Liu-qing Ye ◽  
Zheng-yin Ye ◽  
Kun Ye ◽  
Jie Wu ◽  
Sheng-jie Miao
Keyword(s):  
Drag Coefficient ◽  
Mass Flow ◽  
Design Method ◽  
Leading Edge ◽  
Wave Drag ◽  
Sonic Boom ◽  
Suction Surface ◽  
Trailing Edge ◽  
Flux Flow ◽  
Supersonic Aircraft

AbstractSonic boom reduction has been an urgent need for the development of future supersonic transport, because of the heavy damage of noise pollution. This paper provides a novel concept for supersonic aircraft to reduce the sonic boom and drag coefficient, wherein a suction slot near the leading edge and an injection slot near the trailing edge on the airfoil suction surface are opened. To make sure of a zero net mass flux flow control, the mass flow sucked in near the leading edge is equal to the mass flow injected near the trailing edge. The diamond and NACA0008 airfoils are adopted as the baseline airfoil to verify the capability of the proposed design method. The effects of the suction and injection location, the suction and injection slot size, the mass flow rate and the attack angle on the ground boom signature and drag coefficient are studied in detail. The results show that the optimized airfoils with the suction and injection have benefits in both sonic boom reduction and wave drag reduction. And the reduction of the sonic boom intensity is more sensitive to the injection near the trailing edge than the suction near the leading edge. From the viewpoint of aerodynamics, opening the suction and injection slots will have no adverse effect on the aerodynamic performances of the supersonic aircraft and even increase the lift-drag ratio under some circumstances. For energy saving, the suction and injection slots can be selectively opened, which are opened when the supersonic aircraft flies over the city but are closed when the aircraft flies over the sea.

scholarly journals An active control method for reducing sonic boom of supersonic aircraft

2021 ◽  
Vol 39 (3) ◽  
pp. 566-575
Author(s):  
Liuqing Ye ◽  
Zhengyin Ye ◽  
Boping Ma
Keyword(s):  
Mass Flow Rate ◽  
Active Control ◽  
Mass Flow ◽  
Flow Rate ◽  
Control Method ◽  
Leading Edge ◽  
Sonic Boom ◽  
Trailing Edge ◽  
Supersonic Aircraft ◽  
Active Control Method

Sonic boom reduction has been an urgent need to develop the future supersonic transport, because of the heavy damages of the noise pollution. This paper provides an active control method for the supersonic aircraft to reduce the sonic boom, wherein a suction slot near the leading edge and an injection slot near the trailing edge on the airfoil suction surface are opened, and the mass flow sucked in near the leading edge is equal to the mass flow injected near the trailing edge. The diamond and 566 airfoils are adopted as the baseline airfoil to verify the capability of the active control method, and the effects of the suction and injection location, the mass flow rate and the attack angle on the ground boom signature, the maximum overpressure, the drag coefficients and the ratio of lift to drag are studied in detail. The results show that the proposed active control method can significantly reduce the sonic boom, and the reduction of the sonic boom intensity is more sensitive to the injection near the trailing edge than the suction near the leading edge. Applying this active control method to the diamond (NACA0008) airfoil, when the mass flow rate is 6.5 kg/s(7.5 kg/s), the value of maximum positive overpressure is decreased by 12.87%(12.85%), the value of maximum negative overpressure is decreased by 33.83%(56.77%) and the drag coefficient is decreased by 9.50%(10.96%). It can be seen that the method proposed in this paper has great benefits in the reduction of sonic boom and provides a useful reference for designing a new generation of lower sonic boom supersonic aircraft.

Study on Application of CAE in a Centrifugal Compressor Impeller

2013 ◽  
Vol 787 ◽  
pp. 594-599 ◽  
Author(s):  
Bao Long Gong ◽  
Xiu Jie Jia ◽  
Guang Cun Wang ◽  
Zi Wu Liu
Keyword(s):  
Centrifugal Compressor ◽  
Equivalent Stress ◽  
Leading Edge ◽  
Middle Part ◽  
Turbulent Intensity ◽  
Erosion Wear ◽  
Suction Surface ◽  
Trailing Edge ◽  
Wear Area ◽  
Pressure Surface

CAE technology is the most common method to study properties of impeller in a centrifugal compressor. The fluid field was numerically simulated by CFX program to obtain the distribution rules of pressure, turbulence intensity and erosion wear. Based on fluid-solid interaction, stress and deformation were analyzed by Ansys program. According to the simulation results, the maximum deformation and equivalent stress of the impeller are all located on the junction between the blade trailing edge and the shroud. The most serious damaged part by erosion wear in impeller is on the pressure surface of long blade. The erosion wear area on blade pressure surface caused by particles impact primarily locates on blade trailing edge root and middle part. In the flow field, the turbulent intensity on suction surface is greater than that on pressure surface in the corresponding position and the greatest turbulent intensity is located on the leading edge of suction surface. There is backflow phenomenon around the suction surface of long blade and the short blade has significant effect to reduce backflow. The results of numerical simulation explain some actual impeller failure cases and can be applied to anti-wear impeller design and repair.

Numerical Investigation of the Effect of Inlet Flow Angle on Film Cooling Performance for the Blade Tip With Suction Surface Squealer Rim

2019 ◽  
Author(s):  
Zhiqiang Yu ◽  
Jianjun Liu ◽  
Chen Li ◽  
Baitao An
Keyword(s):  
Mass Flow ◽  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Flow Ratio ◽  
Suction Surface ◽  
Cooling Performance ◽  
Tip Leakage ◽  
Pressure Surface ◽  
Mass Flow Ratio

Abstract Numerical investigations have been performed to study the effect of incidence angle on the aerodynamic and film cooling performance for the suction surface squealer tip with different film-hole arrangements at τ = 1.5% and BR = 1.0. Meanwhile, the full squealer tip as baseline is also investigated. Three incidence angles at design condition (0 deg) and off-design conditions (± 7 deg) are investigated. The suction surface, pressure surface, and the camber line have seven holes each, with an extra hole right at the leading edge. The Mach number at the cascade inlet and outlet are 0.24 and 0.52, respectively. The results show that the incidence angle has a significant effect on the tip leakage flow characteristics and coolant flow direction. The film cooling effectiveness distribution is altered, especially for the film holes near the leading edge. When the incidence angle changes from +7 deg to 0 and −7 deg, the ‘re-attachment line’ moves downstream and the total tip leakage mass flow ratio decreases, but the suction surface tip leakage mass flow ratio near leading edge increases. In general, the total tip leakage mass flow ratio for suction surface squealer tip is 1% greater than that for full squealer tip at the same incidence angle. The total pressure loss coefficient of suction surface squealer tip is larger than that for full squealer tip. The full squealer tip with film holes near suction surface and the suction surface squealer tip with film hole along camber line show high film cooling performance, and the area averaged film cooling effectiveness at positive incidence angle +7 deg is higher than that at 0 and −7 deg. The coolant discharged from film holes near pressure surface only cools narrow region near pressure surface.

An experience-independent inverse design optimization method of compressor cascade airfoil

2019 ◽  
Vol 233 (4) ◽  
pp. 431-442 ◽  
Author(s):  
Yujie Zhu ◽  
Yaping Ju ◽  
Chuhua Zhang
Keyword(s):  
Design Optimization ◽  
Design Method ◽  
Three Dimensional ◽  
Optimal Solution ◽  
Leading Edge ◽  
Optimization Method ◽  
Inverse Design ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Inverse Design Method

Most of the inverse design methods of turbomachinery experience the shortcoming where the target aerodynamic parameters need to be manually specified depending on the designers’ experience and insight, making the design result aleatory and even deviated from the real optimal solution. To tackle this problem, an experience-independent inverse design optimization method is proposed and applied to the redesign of a compressor cascade airfoil in this study. The experience-independent inverse design optimization method can automatically obtain the target pressure distribution along the cascade airfoil through the genetic algorithm, rather than through the manual specification approach. The shape of cascade airfoil is then solved by the adjoint method. The effectiveness of the experience-independent inverse design optimization method is demonstrated by two inverse design cases of the compressor cascade airfoil, i.e. the inverse design of only the suction surface and the inverse design of both the suction and pressure surfaces. The results show that the proposed inverse design method is capable of significantly improving the aerodynamic performance of the compressor cascade. At the examined flow condition, a thin airfoil profile is beneficial to flow accelerations near the leading edge and flow separation avoidance near the trailing edge. The proposed inverse design method is quite generic and can be extended to the three-dimensional inverse design of advanced compressor blades.

Experimental Investigation of Innovative Cooling Schemes on an Additively Manufactured Engine Scale Turbine Nozzle Guide Vane

2020 ◽  
Author(s):  
Mohammad A. Hossain ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Jet Impingement ◽  
Development Effort ◽  
Suction Surface ◽  
Trailing Edge ◽  
Pin Fin ◽  
Nozzle Guide ◽  
Film Cooling Holes

Abstract This study includes the design, validation, and fabrication via Direct Metal Laser Sintering (DMLS) of a gas turbine nozzle guide vanes (NGV) that incorporates three innovative cooling schemes specifically enabled by additive manufacturing. The novel NGV design is the culmination of an extensive research and development effort over a period of four years that included low and high speed cascade testing coupled with unsteady CFD for numerous candidate innovative cooling architectures. The final vane design (SJ-vane) consists of sweeping jet (SJ) film cooling holes on the suction surface, sweeping jet impingement holes at the leading edge and double-wall partial length triangular pin-fin with impinging jet at the trailing edge. For comparison purposes, a second DMLS enabled vane (777-vane) was designed and fabricated with prototypical cooling circuits to serve as a baseline. This vane consists of a shaped film cooling holes on the suction surface, circular impingement holes at the leading edge and full length cylindrical pin-fins at the trailing edge. Experiments with the two DMLS enabled vanes were performed at the Ohio State University Turbine Reacting Flow Rig (TuRFR) at engine relevant temperature (1375K) and Mach number conditions. Infrared (IR) thermography was utilized to measure the wall temperature of the pressure and suction surface at several coolant mass flow rates to estimate the overall cooling effectiveness (ϕ). Results showed improved cooling performance for the advanced cooling schemes (sweeping jet film cooling, impingement cooling and triangular pin-fin cooling) compared to the baseline cooling schemes.

The Limit Inlet Temperature of Refrigeration Centrifugal Compressor to Avoid Phase Change

2020 ◽  
Author(s):  
Duan YaFei ◽  
Tang YongHong ◽  
Jin ZhiHong ◽  
Zou HanSen ◽  
Xi Guang
Keyword(s):  
Phase Change ◽  
Mass Flow ◽  
Centrifugal Compressor ◽  
Leading Edge ◽  
Absolute Velocity ◽  
Inlet Temperature ◽  
Suction Surface ◽  
Pressure Surface ◽  
User Friendly ◽  
Liquid Formation

Abstract From the polytropic compression work formula, we can find that the consumed polytropic work will reduce with the decrease of inlet temperature while compressing the refrigerant to the same compression ratio. However, the refrigerant may condense if the inlet temperature is low enough. Though the principle that the acceleration of fluid may result in condensation has been proved by numerical simulations and experiments, and the liquid formation inside the supercritical carbon dioxide (SCO2) centrifugal compressor has been widely studied, there is still not a user-friendly method to predict whether the inlet condition may cause liquid formation inside the compressor. The fluid flow in the space near the blade suction face of the leading edge (SNSL) is assumed to the similar flow in a converging nozzle when the mass flow is larger enough; the fluid impinges on the suction surface of blades, and the absolute velocity of fluid will not be greater than sound velocity. The fluid turns to impinge on the pressure surface with the decrease of mass flow rate, which is similar to the flow in a converging-diverging nozzle, and the maximum absolute velocity in the SNSL may be greater than the sound speed. A method is proposed to predict the lowest inlet temperature of refrigeration centrifugal compressor to avoid phase change, which is called the limit temperature. The predicted lowest temperature shares the same trend with the numerical results. The condensation will occur inside the compressor when the inlet temperature is lower than the limit inlet temperature. The lowest temperature will first increase and then decrease as the mass flow increases, which should be taken into account while designing a refrigeration centrifugal compressor or adjusting the operating condition.

Vortex structure analysis of unsteady cloud cavitating flows around a hydrofoil

2016 ◽  
Vol 30 (02) ◽  
pp. 1550275 ◽  
Author(s):  
Yu Zhao ◽  
Guoyu Wang ◽  
Biao Huang
Keyword(s):  
Vortex Structure ◽  
Cavity Growth ◽  
Leading Edge ◽  
Lower Surface ◽  
Cavitating Flows ◽  
Suction Surface ◽  
Trailing Edge ◽  
Numerical Predictions ◽  
Moderate Reynolds Number ◽  
Edge Vortex

In this paper, time dependent vortex structures are numerically analyzed for both noncavitating and cloud cavitating flows around a Clark-Y hydrofoil with angle of attack [Formula: see text] at a moderate Reynolds number, [Formula: see text]. The numerical simulations are performed using a transport equation-based cavitation model and the large eddy simulation (LES) approach with a classical eddy viscosity subgrid scale (SGS) model. Compared with experimental results, present numerical predictions are capable of capturing the initiation of cavity, growth toward the trailing edge and subsequent shedding process. Results indicate that in noncavitating conditions, the trailing edge vortex and induced positive vortex shed periodically into the wake region to form the vortex street. In cloud cavitating conditions, interrelations between cavity and vortex induce different vortex dynamics at different cavity developing stages. (i) As attached cavity grows, vorticity production is greatly enhanced by the favorable pressure gradient at the leading edge. The trailing edge flow does not have a direct impact on the attached cavity expansion process. Furthermore, the liquid–vapor interface that moves toward the trailing edge enhances the vorticity in the attached cavity closure region. (ii) When the stable attached sheet cavity grows to its maximum length, the accumulation process of vorticity is eventually interrupted by the formation of the re-entrant jet. Re-entrant jet’s moving upstream leads to a higher spreading rate of the attached cavity and the formation of a large coherent structure inside the attached cavity. Moreover, the wavy/bubbly cavity interface enhances the vorticity near the trailing edge. (iii) As the attached sheet cavity breaks up, this large vortex structure converts toward the trailing edge region, which will eventually couple with a trailing edge vortex shedding from the lower surface to form the cloud cavity. The breakup of the stable attached cavity is the main reason for the vorticity enhancement near the suction surface.

Aeroelastic Instability in Transonic Fans

2015 ◽  
Vol 138 (2) ◽  
Author(s):  
Mehdi Vahdati ◽  
Nick Cumpsty
Keyword(s):  
Mass Flow ◽  
Three Dimensional ◽  
Pressure Rise ◽  
Leading Edge ◽  
Operating Conditions ◽  
Mean Flow ◽  
Suction Surface ◽  
Blade Vibration ◽  
Stall Flutter ◽  
Fan Blade

This paper describes stall flutter, which can occur at part speed operating conditions near the stall boundary. Although it is called stall flutter, this phenomenon does not require the stalling of the fan blade in the sense that it can occur when the slope of the pressure rise characteristic is still negative. This type of flutter occurs with low nodal diameter forward traveling waves and it occurs for the first flap (1F) mode of blade vibration. For this paper, a computational fluid dynamics (CFD) code has been applied to a real fan of contemporary design; the code has been found to be reliable in predicting mean flow and aeroelastic behavior. When the mass flow is reduced, the flow becomes unstable, resulting in flutter or in stall (the stall perhaps leading to surge). When the relative tip speed into the fan rotor is close to sonic, it is found (by measurement and by computation) that the instability for the fan blade considered in this work results in flutter. The CFD has been used like an experimental technique, varying parameters to understand what controls the instability behavior. It is found that the flutter for this fan requires a separated region on the suction surface. It is also found that the acoustic pressure field associated with the blade vibration must be cut-on upstream of the rotor and cut-off downstream of the rotor if flutter instability is to occur. The difference in cut off conditions upstream and downstream is largely produced by the mean swirl velocity introduced by the fan rotor in imparting work and pressure rise to the air. The conditions for instability therefore require a three-dimensional geometric description and blades with finite mean loading. The third parameter that governs the flutter stability of the blade is the ratio of the twisting motion to the plunging motion of the 1F mode shape, which determines the ratio of leading edge (LE) displacement to the trailing edge (TE) displacement. It will be shown that as this ratio increases the onset of flutter moves to a lower mass flow.

Effect of Surface Roughness Location and Reynolds Number on Compressor Cascade Performance

2010 ◽  
Author(s):  
Seung Chul Back ◽  
Garth V. Hobson ◽  
Seung Jin Song ◽  
Knox T. Millsaps
Keyword(s):  
Surface Roughness ◽  
Reynolds Number ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Trailing Edge ◽  
Blade Loading ◽  
Linear Compressor Cascade

An experimental investigation has been conducted to characterize the influence of surface roughness location and Reynolds number on compressor cascade performance. Flow field surveys have been conducted in a low-speed, linear compressor cascade. Pressure, velocity, and flow angles have been measured via a 5-hole probe, pitot probe, and pressure taps on the blades. In addition to the entirely smooth and entirely rough blade cases, blades with roughness covering the leading edge; pressure side; and 5%, 20%, 35%, 50%, and 100% of suction side from the leading edge have been studied. All of the tests have been done for Reynolds number ranging from 300,000 to 640,000.Cascade performance (i.e. blade loading, loss, and deviation) is more sensitive to roughness on the suction side than pressure side. Roughness near the trailing edge of suction side increases loss more than that near the leading edge. When the suction side roughness is located closer to the trailing edge, the deviation and loss increase more rapidly with Reynolds number. For a given roughness location, there exists a Reynolds number at which loss begins to visibly increase. Finally, increasing the area of rough suction surface from the leading edge reduces the Reynolds number at which the loss coefficient begins to increase.

Experimental Investigation of Innovative Cooling Schemes on an Additively Manufactured Engine Scale Turbine Nozzle Guide Vane

2021 ◽  
pp. 1-39
Author(s):  
Mohammad A. Hossain ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey Bons
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Jet Impingement ◽  
Development Effort ◽  
Suction Surface ◽  
Trailing Edge ◽  
Pin Fin ◽  
Nozzle Guide ◽  
Film Cooling Holes

Abstract This study includes the design, validation, and fabrication via Direct Metal Laser Sintering (DMLS) of a gas turbine nozzle guide vanes (NGV) that incorporates three innovative cooling schemes specifically enabled by additive manufacturing. The novel NGV design is the culmination of an extensive research and development effort over a period of four years that included low and high speed cascade testing coupled with unsteady CFD for numerous candidate innovative cooling architectures. The final vane design (SJ-vane) consists of sweeping jet (SJ) film cooling holes on the suction surface, sweeping jet impingement holes at the leading edge and double-wall partial length triangular pin-fin with impinging jet at the trailing edge. For comparison purposes, a second DMLS enabled vane (777-vane) was designed and fabricated with prototypical cooling circuits to serve as a baseline. This vane consists of a shaped film cooling holes on the suction surface, circular impingement holes at the leading edge and full length cylindrical pin-fins at the trailing edge. Experiments with the two DMLS enabled vanes were performed at the Ohio State University Turbine Reacting Flow Rig (TuRFR) at engine relevant temperature (1375 K) and Mach number conditions. Infrared (IR) thermography was utilized to measure the wall temperature of the pressure and suction surface at several coolant mass flow rates to estimate the overall cooling effectiveness (ϕ). Results showed improved cooling performance for the advanced cooling schemes (sweeping jet film cooling, impingement cooling and triangular pin-fin cooling) compared to the baseline cooling schemes.

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