Development, Analysis, and Validation of a Simultaneous Inlet Total Pressure and Swirl Distortion Generator

2020 ◽  
Author(s):  
Dustin J. Frohnapfel ◽  
K. Todd Lowe ◽  
Walter F. O’Brien
Keyword(s):  
Total Pressure ◽  
Experimental Validation ◽  
Computational Analysis ◽  
Pressure Recovery ◽  
Turbofan Engine ◽  
Full Field ◽  
Inlet Distortion ◽  
Swirl Angle ◽  
Total Pressure Recovery ◽  
Ground Test

Abstract Over the last decade, the Turbomachinery and Propulsion Research Laboratory at Virginia Tech has researched, invented, developed, computationally analyzed, experimentally tested, and improved turbofan engine inlet distortion generators. This effort began with modernizing and improving inlet total pressure distortion screens originally conceived over half a century ago; continued with the invention of inlet swirl distortion generators (StreamVanes™) made possible only through advances in modern additive manufacturing technology; and has, thus far, culminated in a novel combined device (ScreenVanes™) capable of simulating realistic flight conditions of coupled inlet total pressure and swirl distortion in a ground-test turbofan engine research platform. The present research focuses on the methodology development, computational analysis, and experimental validation of a novel simultaneous inlet total pressure and swirl distortion generator. A case study involving a single bend S-duct inlet distortion profile demonstrates the ability to generate a high-fidelity profile simulation, yet outlines a design process sufficiently generic for application to any arbitrary inlet geometry or distortion profile. A computational fluid dynamics simulation of the S-duct inlet provided the target profile extracted at the aerodynamic interface plane. Next, utilizing a method of inverse propagation, the planar distortion profile was propagated upstream to yield a flow field that could be manufactured by a distortion generator adequately isolated from turbomachinery effects. The total pressure distortion screen and swirl distortion StreamVane components were then designed and computationally analyzed. Upon successful computational reproduction of the S-duct inlet distortion profile, experimental hardware was fabricated and tested to validate the ScreenVane methodology and distortion generating device. Comparison of the S-duct manufactured distortion and the ScreenVane manufactured distortion was used as the primary criterion for profile replication success. Results from a computational analysis of both the S-duct and ScreenVane indicated excellent agreement in distortion pattern shape, extent, and intensity with full-field total pressure recovery and swirl angle profiles matching within approximately 0.80% and 2.6°, respectively. Furthermore, experimental validation of the ScreenVane indicated nearly identical full-field total pressure recovery and swirl angle profile replication of approximately 1.10% and 2.6°, respectively, when compared to the computational results. The investigation concluded that not only was the ScreenVane device capable of accurately simulating a complex inlet distortion profile, but also produced a viable device for full-scale turbofan engine ground test.

scholarly journals Influence of Propeller Slipstream on the Flow Field of S-Shaped Intake

2021 ◽  
Vol 2021 ◽  
pp. 1-14
Author(s):  
Shuili Ren ◽  
Peiqing Liu
Keyword(s):  
Flow Field ◽  
Total Pressure ◽  
Engine Performance ◽  
Pressure Recovery ◽  
Navier Stokes ◽  
Navier Stokes Equation ◽  
Recovery Coefficient ◽  
Distortion Coefficient ◽  
Total Pressure Recovery ◽  
Pressure Recovery Coefficient

For turboprop engine, the S-shaped intake affects the engine performance and the propeller is not far in front of the inlet of the S-shaped intake, so the slipstream inevitably affects the flow field in the S-shaped intake and the engine performance. Here, an S-shaped intake with/without propeller is studied by solving Reynolds-averaged Navier-Stokes equation employed SST k-ω turbulence model. The results are presented as time-averaged results and transient results. By comparing the flow field in S-shaped intake with/without propeller, the transient results show that total pressure recovery coefficient and distortion coefficient on the AIP section vary periodically with time. The time-averaged results show that the influence of propeller slipstream on the performance of S-shaped intake is mainly circumferential interference and streamwise interference. Circumferential interference mainly affects the secondary flow in the S-shaped intake and then affects the airflow uniformity; the streamwise interference mainly affects the streamwise flow separation in the S-shaped intake and then affects the total pressure recovery. The total pressure recovery coefficient on the AIP section for the S-shaped intake with propeller is 1%-2.5% higher than that for S-shaped intake without propeller, and the total pressure distortion coefficient on the AIP section for the S-shaped intake with propeller is 1%-12% higher than that for the S-shaped intake without propeller. However, compared with the free stream flow velocity ( Ma = 0.527 ), the influence of the propeller slipstream belongs to the category of small disturbance, which is acceptable for engineering applications.

Flow field and performance analysis of an integrated diverterless supersonic inlet

2011 ◽  
Vol 115 (1170) ◽  
pp. 471-480 ◽  
Author(s):  
J. Masud
Keyword(s):  
Mass Flow ◽  
Total Pressure ◽  
Critical Mass ◽  
Wind Tunnel Test ◽  
Flow Behaviour ◽  
Pressure Recovery ◽  
Duct Flow ◽  
Supersonic Inlet ◽  
Total Pressure Recovery ◽  
And Performance

Abstract In this paper the computed flow and performance characteristics at low angle-of-attack (AOA) of an integrated diverterless supersonic inlet (DSI) are presented. The subsonic characteristics are evaluated at M∞ = 0·8 while the supersonic characteristics are evaluated at M∞= 1·7, which is near the design Mach number for the intake. In addition to the external flow features, the internal intake duct flow behaviour is also evaluated. The results of this study indicate effective boundary layer diversion due to the ‘bump’ compression surface in both subsonic and supersonic regimes. At M∞ = 1·7, the shockwave structure (oblique/normal shockwave) on the ‘bump’ compression surface and intake inlet is satisfactory at design (critical) mass flow rate. The intake duct flow behaviour at subsonic and supersonic conditions is generally consistent with ‘Y’ shaped intake duct of the present configuration. The secondary flow structure inside the duct has been effectively captured by present computations. The computed intake total pressure recovery at M∞ = 1·7 exhibits higher-than-conventional behaviour at low mass flow ratios, which is attributed to unique inlet design. Overall computed subsonic and supersonic total pressure recovery characteristics are satisfactory under the evaluated conditions and are also in agreement with wind tunnel test data.

Numerical Study on the Plasma Generated by Different Discharge Modes at the Upstream of Fuel Jet in Scramjet Combustor

2013 ◽  
Vol 444-445 ◽  
pp. 1345-1349
Author(s):  
Si Yin Zhou ◽  
Wan Sheng Nie ◽  
Bo He ◽  
Xue Ke Che ◽  
Xue Min Tian
Keyword(s):  
Total Pressure ◽  
Recirculation Zone ◽  
Pressure Recovery ◽  
Flow Conditions ◽  
Recovery Coefficient ◽  
Discharge Mode ◽  
Scramjet Combustor ◽  
Fuel Jet ◽  
Total Pressure Recovery ◽  
Pressure Recovery Coefficient

How to enhance the combustion and reduce the total pressure loss in scramjet combustor are very critical for the practical application of hypersonic aircraft. Based on the dominant thermal mechanism of arc plasma, the plasma generated in combustor is regarded as a promising method to improve the combustion. As a result, the combustor model with transverse fuel jet and plasma generated by two discharge modes at the upstream of flameholding cavity is established and it is used to study the mechanism of fuel mixing enhancement through numerical investigation. The results show that an oblique shock wave would be formed at the upstream of the pseudo small plasma hump, and interact with the separation shock wave induced by the transverse jet. This results in the recirculation zone at the upstream of fuel jet being enlarged obviously. Besides that, under the non-reaction flow conditions, the total pressure recovery coefficient increases due to the plasma generated. However, the total pressure recovery coefficient varies apparently and the shear layer above the cavity is fluctuant when the plasma is generated by periodical discharge mode. While under the reaction flow conditions, the shear layer develops thicker and the total pressure recovery coefficient decreases. And due to the existing of plasma, the mole fraction of product water increases. But compared with the steady discharge mode, the level of water is lower and the total pressure recovery coefficient decreases more under the periodical discharge mode. Though the plasma generated by steady discharge mode shows better performance in assisting combustion and reducing the pressure loss, considering the energy saving and the use of different parameters of the periodical discharge, the same effects of enhancing the fuel mixing through enlarging the recirculation zone located at the upstream of fuel jet and promoting the mass exchange of cavity can be reached. More numerical experiments have to be done to optimize the parameters of periodical discharge plasma to receive a best improvement on the performance of scramjet combustor.

Numerical analysis on the conceptual design of an air-breathing engine inlet working in mach number 3∼5.5

2021 ◽  
pp. 095441002110113
Author(s):  
Chao Huo ◽  
Zhenhua Yang ◽  
Zhengze Zhang ◽  
Peijin Liu
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Back Pressure ◽  
Pressure Recovery ◽  
Shock Train ◽  
Recovery Coefficient ◽  
Air Breathing ◽  
Supersonic Inlet ◽  
Total Pressure Recovery ◽  
Pressure Recovery Coefficient

Based on the equal-intensity shock theory, this article designed a supersonic inlet working in Mach number 3.0∼5.5 with the background of an air-breathing engine. The inlet applied the four-shock train mixed compression configuration and inserted a sidewall compression at the beginning of the isolator. Through developing effective 3D RANS computations validated by current experiments, the performance of the designed inlet was identified. The designed inlet self-starts at freestream Mach number Ma∞ = 3.0 under which the total pressure recovery coefficient has dramatic increment, and the aerodynamic choking at the inlet throat no longer presents; the inlet keeps working at all studied flight states with zero angle of attack (AoA) and achieves shock-on-lip at the design point Ma∞ = 5.0. Both positive and negative AoAs can disrupt the equal-intensity shock allocations, which degrade the inlet performance. The inlet obtains maximum total pressure recovery coefficient at zero AoA. The maximum back pressure at Ma∞ = 3.0∼5.5 obtained by the inlet surpasses the requirements and keeps a certain margin. The inlet performance basically meets all the goals proposed by the engine.

Strut Losses in a Diverging Annular Diffuser With Swirling Flow

2006 ◽  
Author(s):  
G. K. Feldcamp ◽  
A. M. Birk
Keyword(s):  
Total Pressure ◽  
Swirling Flow ◽  
Pressure Recovery ◽  
Swirl Number ◽  
Pressure Losses ◽  
Swirl Angle ◽  
Wide Range ◽  
Annular Diffuser ◽  
Inlet Swirl ◽  
The Ideal

An experimental investigation into the overall influence of struts spanning a double divergent annular diffuser followed by a straight cored annular diffuser has been undertaken in order to determine the performance of various strut configurations over a wide range of inlet swirl conditions. Two strut profiles have been investigated in four and eight strut configurations. Results have shown that the presence of struts under no swirl conditions have a relatively small effect on the overall total pressure loss. Increasing the inlet swirl angle to 20° has shown that the struts are able to assist in recovery of the swirling flow such that the pressure recovery nearly approaches that without struts, despite increased total pressure losses. Performance at 40° swirl is highly dependent on strut profile; the higher thickness-to-chord ratio strut configurations show minimal decrease in pressure recovery compared to 20° swirl, while the lower thickness-to-chord ratio configurations experiences a significant decrease as the result of significant flow separation from the struts. The exit swirl number has been shown to correlate strongly with the strut profile shape, while the number of struts had only a secondary influence. The exit velocity profiles show significant distortions at 40° swirl, and as a result the ideal pressure recovery calculated from the inlet and exit profiles change with strut configuration at 40° swirl.

Total Pressure Recovery of Flared Fan Nozzles Used as Inlets

1979 ◽  
Vol 16 (2) ◽  
pp. 110-115 ◽  
Author(s):  
T.G. Keith ◽  
T.N. Obee ◽  
D.A. Dietrich
Keyword(s):  
Total Pressure ◽  
Pressure Recovery ◽  
Total Pressure Recovery
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