The Cause of Tone Generation by Aero-Engine Fans at High Subsonic Tip Speeds and the Effect of Forward Speed

1974 ◽  
Vol 96 (3) ◽  
pp. 228-234 ◽  
Author(s):  
N. A. Cumpsty ◽  
B. W. Lowrie
Keyword(s):  
Boundary Layer ◽  
Forward Speed ◽  
Flow Distortion ◽  
Forward Motion ◽  
Inlet Guide Vanes ◽  
Wall Boundary Layer ◽  
Aero Engine ◽  
Mach Numbers ◽  
Static Conditions ◽  
Tone Generation

Modern aero-engine fans (without inlet guide vanes) produce a tone at blade passing frequency by interacting with the inlet flow distortion. Because the distortion is altered by forward motion of the aircraft the tone is also different under static and inflight conditions. At low tip relative Mach numbers (less than about 0.85) the significant source of distortion is the intake wall boundary layer. At higher tip Mach numbers the principal distortion under static conditions is the atmospheric turbulence drawn into the intake, but in flight this reduces and other forms of distortion may then dominate.

Wall Boundary Layers in Turbomachines

1972 ◽  
Vol 14 (6) ◽  
pp. 411-423 ◽  
Author(s):  
H. Marsh ◽  
J. H. Horlock
Keyword(s):  
Boundary Layer ◽  
Integral Equations ◽  
Cross Flow ◽  
Inlet Guide Vanes ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
Alternative Approach ◽  
Flow Through ◽  
The Many ◽  
Momentum Integral

Equations for the passage-averaged flow in a cascade are used to derive the momentum integral equations governing the development of the wall boundary layer in turbomachines. Several existing methods of analysis are discussed and an alternative approach is given which is based on the passage-averaged momentum integral equations. The analysis leads to an anomaly in the prediction of the cross flow and to avoid this it is suggested that for the many-bladed cascade there should be a variation of the blade force through the boundary layer. This variation of the blade force can be included in the analysis as a force deficit integral. The growth of the wall boundary layer has been calculated by four methods and the predictions are compared with two sets of published experimental results for flow through inlet guide vanes.

scholarly journals Numerical Modelling of Shock Wave Boundary Layer Interactions in Aero-Engine Intakes at Incidence

2018 ◽  
Author(s):  
Hardeep S. Kalsi ◽  
Paul G. Tucker
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Computational Cost ◽  
Low Frequency ◽  
Flow Distortion ◽  
Significant Saving ◽  
Wave Boundary Layer ◽  
Aero Engine ◽  
Downstream Flow ◽  
Wave Boundary

During situations of high incidence, high curvature of aero-engine intake lips can locally accelerate flow to supersonic speeds, producing undesirable shock wave boundary layer interactions (SWBLIs). The present work describes simulations of a novel experimental model resembling a lower intake lip at incidence. RANS, LES and hybrid RANS-LES are carried out at two angles of attack, α = 23° and α = 25°, with α = 25° possessing a high degree of shock oscillation. Modifications to the Spalart-Allmaras (SA) RANS turbulence model are proposed to account for re-laminarisation and curvature. These provide an improvement in prediction compared standard SA model. However, RANS models fail to reproduce post shock interaction flow, giving incorrect shape of the flow distortion. LES and hybrid RANS-LES perform well here, with downstream flow distortion in very good agreement with experimental measurements. LES and hybrid RANS-LES also capture the time averaged smearing of the shock which RANS cannot. However, low frequency shock oscillations in the α = 25° case are costly for LES, requiring long simulation time to obtain time averaged flow statistics. Hybrid RANS-LES offers a significant saving in computational cost, costing approximately 20% of LES.

Flow Distortion and Noise Produced by a Thrusting Rotor Ingesting a Planar Turbulent Boundary Layer

2015 ◽  
Author(s):  
David M. Wisda ◽  
Henry Murray ◽  
William N. Alexander ◽  
Michael Nelson ◽  
William J. Devenport ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Turbulent Boundary Layer ◽  
Flow Distortion

Experiments on Flame Flashback in a Quasi-2D Turbulent Wall Boundary Layer for Premixed Methane-Hydrogen-Air Mixtures

2010 ◽  
Author(s):  
Christian Eichler ◽  
Thomas Sattelmayer
Keyword(s):  
Boundary Layer ◽  
Boundary Layers ◽  
Turbulent Flows ◽  
Turbulent Transport ◽  
Wall Friction ◽  
Evaluation Procedure ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
Flame Flashback ◽  
Turbulent Wall

Premixed combustion of hydrogen-rich mixtures involves the risk of flame flashback through wall boundary layers. For laminar flow conditions, the flashback mechanism is well understood and is usually correlated by a critical velocity gradient at the wall. Turbulent transport inside the boundary layer considerably increases the flashback propensity. Only tube burner setups have been investigated in the past and thus turbulent flashback limits were only derived for a fully-developed Blasius wall friction profile. For turbulent flows, details of the flame propagation in proximity to the wall remain unclear. This paper presents results from a new experimental combustion rig, apt for detailed optical investigations of flame flashbacks in a turbulent wall boundary layer developing on a flat plate and being subject to an adjustable pressure gradient. Turbulent flashback limits are derived from the observed flame position inside the measurement section. The fuels investigated cover mixtures of methane, hydrogen and air at various mixing ratios. The associated wall friction distributions are determined by RANS computations of the flow inside the measurement section with fully resolved boundary layers. Consequently, the interaction between flame back pressure and incoming flow is not taken into account explicitly, in accordance with the evaluation procedure used for tube burner experiments. The results are compared to literature values and the critical gradient concept is reviewed in light of the new data.

Some Practical Aspects of Kinetic Heating Calculations

1959 ◽  
Vol 63 (587) ◽  
pp. 637-645 ◽  
Author(s):  
C. L. Bore
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Compressible Flow ◽  
Temperature Variations ◽  
Transfer Rates ◽  
Mach Numbers

This paper is concerned primarily with some of the practical difficulties encountered in connection with the prediction of kinetic heating temperatures. Attention will be concentrated upon methods for estimating temperatures and heat transfer rates for practical aircraft designed to fly at Mach numbers up to about five.One factor common to all kinetic heating calculations is the variation of temperature through the thickness of the boundary layer, with consequent variation of viscosity. At Mach numbers above about 3, these temperature variations also lead to considerable variations of other properties of air—which are commonly assumed to remain constant—even in classical compressible flow aerodynamics. These factors complicate the aerodynamic equations.

scholarly journals Investigation of Flows in Rectangular Diffusers With Inlet Flow Distortion

1982 ◽  
Author(s):  
M. M. Al-Mudhafar ◽  
M. Ilyas ◽  
F. S. Bhinder
Keyword(s):  
Experimental Study ◽  
Velocity Profiles ◽  
Pressure Recovery ◽  
Two Dimensional ◽  
Flow Distortion ◽  
Inlet Flow ◽  
Inlet Velocity ◽  
Inlet Distortion ◽  
Mach Numbers ◽  
Inlet Flow Distortion

The results of an experimental study on the influence of severely distorted velocity profiles on the performance of a straight two-dimensional diffuser are reported. The data cover entry Mach numbers ranging from 0.1 to 0.6 and several inlet distortion levels. The pressure recovery progressively deteriorates as the inlet velocity is distorted.

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