A Note on the “Isentropic Ramp” Intake Diffuser in the Mach Number Range 6 to 10

1967 ◽  
Vol 71 (673) ◽  
pp. 49-51
Author(s):  
E. R. Bergstrom
Keyword(s):  
Mach Number ◽  
Order Approximation ◽  
Combustion Process ◽  
Inviscid Flow ◽  
Oblique Shock Wave ◽  
Pressure Gradients ◽  
Number Range ◽  
Hypersonic Boundary Layers ◽  
Layer Flow ◽  
Inlet Conditions

A need exists for empirical criteria by which the behaviour of hypersonic boundary layers flowing in adverse pressure gradients can be predicted, thereby placing the design of high Mach number intake/diffuser systems on a rational basis.The main results of the inviscid flow performance analysis of a two-dimensional “isentropic ramp” intake diffusing to supersonic Mach numbers over the free stream Mach number range 6 to 10 are presented. Limitations on the degree of diffusion, and hence geometry, have been deduced in terms of the combustion length requirements of an idealised hydrogen/air combustion process with com-bustor inlet conditions given by the intake. The analysis can be considered as a first order approximation defining the ranges of continuous pressure gradients and oblique shock wave strengths relevant to the experimental investigation of boundary layer flow in the mixed adverse pressure gradients associated with hypersonic intakes.

Performance of an Annular Combustor Under Windmill Conditions During Stand-Alone and Engine Level Altitude Test

2017 ◽  
Author(s):  
Srinivasan Karuppannan ◽  
Dalton Maurya ◽  
Raju D. Navindgi ◽  
N. Muthuveerappan
Keyword(s):  
Mach Number ◽  
Experimental Test ◽  
Total Pressure ◽  
Development Stage ◽  
Flight Altitude ◽  
Number Range ◽  
Turbine Engine ◽  
Annular Combustor ◽  
Engine Compressor ◽  
Inlet Conditions

Relight envelope of the combustor needs to be experimentally generated and established during the design and development of an aero gas turbine engine. Usually, during development stage of engine, compressor characteristics are not readily available at such low speeds and hence, it becomes difficult to specify the combustor inlet conditions such as pressure, temperature and Mach number during the engine light up studies. This paper compares the experimental test data generated on an annular combustor for windmill conditions during stand-alone mode and engine level tests under simulated flight conditions. The stand-alone combustor trials were conducted for the range of total pressure and temperature relevant to the flight altitude and Mach number range. During the engine level tests, combustor relight tests were conducted under simulated conditions (ISA+15) for altitudes ranging from 5.5 km to 10 km, flight Mach numbers in the range of 0.45 to 0.80. In this paper, effect of altitude and flight Mach number on the windmill spool speed, combustor pressure and temperature are studied.

Performance Margins of Non-Reflecting Slotted Walls in a Transonic Linear Cascade Tunnel

2006 ◽  
Author(s):  
A. Rona ◽  
J. P. Gostelow
Keyword(s):  
Mach Number ◽  
Inviscid Flow ◽  
Number Range ◽  
Linear Cascade ◽  
Low Pressure Turbine ◽  
Trailing Edges ◽  
Flow Phenomena ◽  
Interior Flow ◽  
End Wall ◽  
Appreciable Reduction

An experimental investigation was conducted into the endwall interference from a low-pressure turbine nozzle blade profile tested in linear cascade at an isentropic discharge Mach number of 1.27. This was above the profile design point of Mach 0.995. This highlighted that inviscid flow phenomena, at the cascade pitchwise boundaries, are the main cause of spurious end-wall interference at these test conditions. Specifically, fish-tail shocks from the profile trailing edges reflect at these boundaries disturbing the interior flow. An appreciable reduction of such interference was obtained through the use of a slotted tailboard downstream of the outmost blade trailing edge, with void ratio and pitch optimized to operate at Mach 1.27 in this cascade wind tunnel. Cascade tests in the Mach number range 1.20 to 1.32 show that the tailboard at off-design conditions gives a more pitchwise periodic discharge than without a tailboard and a test section open jet boundary. Reducing the tailboard pitch from its design value of 64° to 62° does not further improve the flow pitchwise periodicity over the Mach number range 1.18 to 1.31. Over this range, the slotted tailboard at its design pitch angle retains an appreciable performance margin over the other two end-wall configurations tested.

Auto and Cross Correlation Measurements in a Turbine Cascade Using a Digital Correlator

1982 ◽  
Author(s):  
P. J. Bryanston-Cross ◽  
J. J. Camus
Keyword(s):  
Mach Number ◽  
High Speed ◽  
Cross Correlation ◽  
Turbine Blades ◽  
Image Plane ◽  
Strongly Correlated ◽  
Trailing Edge ◽  
Number Range ◽  
Cross Correlations ◽  
Digital Correlator

A simple technique has been developed which samples the dynamic image plane information of a schlieren system using a digital correlator. Measurements have been made in the passages and in the wakes of transonic turbine blades in a linear cascade. The wind tunnel runs continuously and has independently variable Reynolds and Mach number. As expected, strongly correlated vortices were found in the wake and trailing edge region at 50 KHz. Although these are strongly coherent we show that there is only limited cross-correlation from wake to wake over a Mach no. range M = 0.5 to 1.25 and variation of Reynolds number from 3 × 105 to 106. The trailing edge fluctuation cross correlations were extended both upstream and downstream and preliminary measurements indicate that this technique can be used to obtain information on wake velocity. The vortex frequency has also been measured over the same Mach number range for two different cascades. The results have been compared with high speed schlieren photographs.

scholarly journals Performance Assessment of an Annular S-Shaped Duct

1995 ◽  
Author(s):  
D. W. Bailey ◽  
K. M. Britchford ◽  
J. F. Carrotte ◽  
S. J. Stevens
Keyword(s):  
Shear Stress ◽  
Static Pressure ◽  
Compressor Blade ◽  
Turbine Engines ◽  
Pressure Gradients ◽  
Gas Turbine Engines ◽  
Entry Length ◽  
Static Pressure Distribution ◽  
Flow Curvature ◽  
Inlet Conditions

An experimental investigation has been carried out to determine the aerodynamic performance of an annular S-shaped duct representative of that used to connect the compressor spools of aircraft gas turbine engines. For inlet conditions in which boundary layers are developed along an upstream entry length the static pressure, shear stress and velocity distributions are presented. The data shows that as a result of flow curvature significant streamwise pressure gradients exist within the duct, with this curvature also affecting the generation and suppression of turbulence. The stagnation pressure loss within the duct is also assessed and is consistent with the measured distributions of shear stress. More engine representative conditions are provided by locating a single stage compressor at inlet to the duct. Relative to the naturally developed inlet conditions the flow within the duct is less likely to separate, but mixing out of the compressor blade wakes increases the measured duct loss. With both types of inlet conditions the effect of a radial strut, such as that used for carrying loads and engine services, is also described both in terms of the static pressure distribution along the strut and its contribution to overall loss.

Mach Number Influence on Reduced-Order Models of Inviscid Potential Flows in Turbomachinery

2002 ◽  
Vol 124 (4) ◽  
pp. 977-987 ◽  
Author(s):  
Bogdan I. Epureanu ◽  
Earl H. Dowell ◽  
Kenneth C. Hall
Keyword(s):  
Mach Number ◽  
Steady Flow ◽  
Inviscid Flow ◽  
Reduced Order Models ◽  
Reduced Order ◽  
Spatial Gradients ◽  
Potential Flows ◽  
Proper Orthogonal Decomposition Technique ◽  
Flow Through ◽  
Phase Angles

An unsteady inviscid flow through a cascade of oscillating airfoils is investigated. An inviscid nonlinear subsonic and transonic model is used to compute the steady flow solution. Then a small amplitude motion of the airfoils about their steady flow configuration is considered. The unsteady flow is linearized about the nonlinear steady response based on the observation that in many practical cases the unsteadiness in the flow has a substantially smaller magnitude than the steady component. Several reduced-order modal models are constructed in the frequency domain using the proper orthogonal decomposition technique. The dependency of the required number of aerodynamic modes in a reduced-order model on the far-field upstream Mach number is investigated. It is shown that the transonic reduced-order models require a larger number of modes than the subsonic models for a similar geometry, range of reduced frequencies and interblade phase angles. The increased number of modes may be due to the increased Mach number per se, or the presence of the strong spatial gradients in the region of the shock. These two possible causes are investigated. Also, the geometry of the cascade is shown to influence strongly the shape of the aerodynamic modes, but only weakly the required dimension of the reduced-order models.

The Stability of a Two-Layer Inviscid Flow Between an Elastic Layer and a Rigid Boundary

1999 ◽  
Vol 66 (1) ◽  
pp. 197-203 ◽  
Author(s):  
P. J. Schall ◽  
J. P. McHugh
Keyword(s):  
Elastic Layer ◽  
Inviscid Flow ◽  
Elastic Parameters ◽  
Rigid Boundary ◽  
Two Fluids ◽  
Coating Flow ◽  
Fluid Layers ◽  
Mode 2 ◽  
Layer Flow ◽  
The Stability

The linear stability of two-layer flow over an infinite elastic substraw is considered. The problem is motivated by coating flow in a printing press. The flow is assumed to be inviscid and irrotational. Surface tension between the fluid layers is included, but gravity is neglected. The results show two unstable modes: one mode associated with the interface between the elastic layer and the fluid (mode 1), and the other concentrated on the interface between the two fluids (mode 2). The behavior of the unstable modes is examined while varying the elastic parameters, and it is found that mode 1 can be made stable, but mode 2 remains unstable at small wavenumber, similar to the classic Kelvin—Helmholtz mode.

scholarly journals Influence of parameters on flame expansion in a high-speed flow: Experimental and numerical study

2019 ◽  
Vol 234 (8) ◽  
pp. 1122-1130 ◽  
Author(s):  
Shilong Zhao ◽  
Fan Yuxin ◽  
Zhang Xiaolei
Keyword(s):  
Image Processing ◽  
High Speed ◽  
Numerical Study ◽  
Combustion Process ◽  
Engineering Practice ◽  
Blockage Ratio ◽  
Lean Premixed Combustion ◽  
Expansion Angle ◽  
Inlet Conditions ◽  
Increasing Temperature

Flameholder-stabilized flames are conventional and also commonly used in propulsion and various power generation fields to maintain combustion process. The characteristics of flame expansion were obtained with various blockage ratios, which were observed to be highly sensitive to inlet conditions such as temperatures and velocities. Experiments and simulations combined methodology was performed; also the approach adopted on image processing was calculated automatically through a program written in MATLAB. It was found that the change of flame expansion angle indicated increasing fuel supply could contribute to the growth of flame expansion angle in lean premixed combustion. Besides, the influence of inlet velocity on flame expansion angle varies with different blockage ratios, i.e. under a small blockage ratio (BR = 0.1), flame expansion angle declined with the increase of velocity; however, under a larger blockage ratio (BR = 0.2 or 0.3), flame expansion angle increased firstly and then decreased with the increasing velocity. Likewise, flame expansion angle increased firstly and then decreased with the increasing temperature under BR = 0.2/0.3. In addition, flame expansion angle was almost the same for BR = 0.2 and BR = 0.3 at a higher temperature (900 K), and both of which were bigger than BR = 0.1. Overall, BR = 0.2 is the best for increasing flame expansion angle and reducing total pressure loss. The influence of velocity and temperature on flame expansion angle found from this research are vital for engineering practice and for developing a further image processing method to measure flame boundary.

A Comparison of the Influence of Fuel/Air Unmixedness on NOx Emissions in Lean Premixed, Non-Catalytic and Catalytically Stabilized Combustion

1997 ◽  
Author(s):  
A. Schlegel ◽  
M. Streichsbier ◽  
R. Mongia ◽  
R. Dibble
Keyword(s):  
Combustion Chamber ◽  
Catalytic Combustion ◽  
Combustion Process ◽  
Tubular Reactor ◽  
Optical Probe ◽  
Nox Emissions ◽  
Fuel Concentration ◽  
Lean Premixed ◽  
Inlet Conditions ◽  
Over Time

Experimental results on the influence of temporal unmixedness on NOx emissions are presented for both non-catalytic and catalytically stabilized, lean premixed combustion. The test rig used for the experiments consists of a fuel/air mixing section which allows variation of the degree of temporal unmixedness while maintaining a uniform “average over time” concentration profile over the cross section at the inlet to the combustion chamber. The unmixedness is measured as “rms fluctuations in fuel concentration” by an optical probe using laser absorption at 3.39μm over a 9mm gap. “Average over time” measurements are taken with “conventional” suction probe analyzers. The combustion chamber is an insulated, tubular reactor (i.d. 26.4mm). At the inlet to the combustion chamber a honeycomb monolith section is inserted. This monolith is either catalytically active or inactive for catalytically stabilized or non-catalytic combustion respectively. For both modes, the exact same inlet conditions are applied. In catalytically stabilized combustion a fraction of the fuel is consumed within the catalyst and the remaining fuel is burnt in the subsequent homogeneous combustion zone. It is shown that catalytically stabilized combustion yields lower NOx emissions and, more important, that the effect of temporal fuel/air unmixedness on NOx emissions is much smaller than with non-catalytic combustion under identical inlet conditions. Experimental evidence leads to the conclusion, that the catalyst is capable of reducing temporal fluctuations in fuel concentration and/or temperature in the combustion process, thereby preventing excess NOx formation. As a result, the requirements on mixing quality are less stringent when using catalytically stabilized combustion instead of conventional, non-catalytic combustion.

scholarly journals Self-similar compressible turbulent boundary layers with pressure gradients. Part 1. Direct numerical simulation and assessment of Morkovin’s hypothesis

2019 ◽  
Vol 880 ◽  
pp. 239-283 ◽  
Author(s):  
Christoph Wenzel ◽  
Tobias Gibis ◽  
Markus Kloker ◽  
Ulrich Rist
Keyword(s):  
Numerical Simulation ◽  
Boundary Layer ◽  
Mach Number ◽  
Direct Numerical Simulation ◽  
Boundary Layers ◽  
Turbulent Boundary Layers ◽  
Shear Stresses ◽  
Pressure Gradients ◽  
Mean Flow ◽  
Self Similar

A direct numerical simulation study of self-similar compressible flat-plate turbulent boundary layers (TBLs) with pressure gradients (PGs) has been performed for inflow Mach numbers of 0.5 and 2.0. All cases are computed with smooth PGs for both favourable and adverse PG distributions (FPG, APG) and thus are akin to experiments using a reflected-wave set-up. The equilibrium character allows for a systematic comparison between sub- and supersonic cases, enabling the isolation of pure PG effects from Mach-number effects and thus an investigation of the validity of common compressibility transformations for compressible PG TBLs. It turned out that the kinematic Rotta–Clauser parameter $\unicode[STIX]{x1D6FD}_{K}$ calculated using the incompressible form of the boundary-layer displacement thickness as length scale is the appropriate similarity parameter to compare both sub- and supersonic cases. Whereas the subsonic APG cases show trends known from incompressible flow, the interpretation of the supersonic PG cases is intricate. Both sub- and supersonic regions exist in the boundary layer, which counteract in their spatial evolution. The boundary-layer thickness $\unicode[STIX]{x1D6FF}_{99}$ and the skin-friction coefficient $c_{f}$, for instance, are therefore in a comparable range for all compressible APG cases. The evaluation of local non-dimensionalized total and turbulent shear stresses shows an almost identical behaviour for both sub- and supersonic cases characterized by similar $\unicode[STIX]{x1D6FD}_{K}$, which indicates the (approximate) validity of Morkovin’s scaling/hypothesis also for compressible PG TBLs. Likewise, the local non-dimensionalized distributions of the mean-flow pressure and the pressure fluctuations are virtually invariant to the local Mach number for same $\unicode[STIX]{x1D6FD}_{K}$-cases. In the inner layer, the van Driest transformation collapses compressible mean-flow data of the streamwise velocity component well into their nearly incompressible counterparts with the same $\unicode[STIX]{x1D6FD}_{K}$. However, noticeable differences can be observed in the wake region of the velocity profiles, depending on the strength of the PG. For both sub- and supersonic cases the recovery factor was found to be significantly decreased by APGs and increased by FPGs, but also to remain virtually constant in regions of approximated equilibrium.

Sign in / Sign up

Export Citation Format

Share Document