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.

Numerical investigation of critical lip thickness of subsonic co-flowing jet

2021 ◽  
Vol ahead-of-print (ahead-of-print) ◽  
Author(s):  
Naren Shankar R. ◽  
Ganesan V.G. ◽  
Dilip Raja N. ◽  
Sathish Kumar K. ◽  
Vijayaraja K.
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Static Pressure ◽  
Velocity Ratio ◽  
Critical Limit ◽  
Number Range ◽  
Content Type ◽  
Jet Mixing ◽  
Pressure Decay ◽  
Mach Numbers

Purpose The effect of increasing lip thickness (LT) and Mach number on subsonic co-flowing Jet (CFJ) decay at subsonic and correctly expanded sonic Mach numbers has been analysed experimentally and numerically in this study. This study aims to a critical LT below which mixing enhances and above which mixing inhibits. Design/methodology/approach LT is the distance, separating the primary nozzle and the secondary duct, present in the co-flowing nozzle. The CFJ with LT ranging from 2 mm to 150 mm at jet exit Mach numbers of 0.6, 0.8 and 1.0 were studied in detail. The CFJ with 2 mm LT is used for comparison. Centreline total pressure decay, centreline static pressure decay and near field flow behaviour were analysed. Findings The result shows that the mixing enhances until a critical limit and a further increase in the LT does not show any variation in the jet mixing. Beyond this critical limit, the secondary jet has a detrimental effect on the primary jet, which deteriorates the process of mixing. The CFJ within the critical limit experiences a significantly higher mixing. The effect of the increase in the Mach number has marginal variation in the total pressure and significant variation in static pressure along the jet axis. Practical implications In this study, the velocity ratio (VR) is maintained constant and the bypass ratio (BR) was varied from low value to very high values for subsonic and correctly expanded sonic. Presently, commercial aircraft engine operates under these Mach numbers and low to ultra-high BR. Hence, the present study becomes essential. Originality/value This is the first effort to find the critical value of LT for a constant VR for a Mach number range of 0.6 to 1.0, compressible CFJ. The CFJs with constant VR of unity and varying LT, in these Mach number range, have not been studied in the past.

Effect of Nonuniform Inlet Conditions on Endwall Secondary Flows

2002 ◽  
Vol 124 (4) ◽  
pp. 623-631 ◽  
Author(s):  
K. S. Hermanson ◽  
K. A. Thole
Keyword(s):  
Secondary Flow ◽  
Total Pressure ◽  
Large Scale ◽  
Secondary Flows ◽  
Turbine Engine ◽  
Thermal Fields ◽  
Pressure Profiles ◽  
Stator Vane ◽  
Inlet Conditions ◽  
Computational Predictions

Exit combustor flow and thermal fields entering downstream stator vane passages in a gas turbine engine are highly nonuniform. These flow and thermal fields can significantly affect the development of the secondary flows in the turbine passages contributing to high platform heat transfer and large aerodynamic losses. The flow and thermal fields combine to give nonuniform total pressure profiles entering the turbine passage which, along with the airfoil geometry, dictate the secondary flow field. This paper presents an analysis of the effects of varying total pressure profiles in both the radial and combined radial and circumferential directions on the secondary flowfields in a first-stage stator vane. These inlet conditions used for the first vane simulations are based on the exit conditions predicted for a combustor. Prior to using the predictions, these CFD simulations were benchmarked against flowfield data measured in a large-scale, linear, turbine vane cascade. Good agreement occurred between the computational predictions and experimentally measured secondary flows. Analyses of the results for several different cases indicate variations in the secondary flow pattern from pitch to pitch, which attributes to the rationale as to why some airfoils quickly degrade while others remain intact over time.

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.

A Ground Test Method and the Error Analysis of the Heat Exchanger in the Intake of Aircraft

2014 ◽  
Vol 1025-1026 ◽  
pp. 137-142
Author(s):  
Yu Fei Cai ◽  
Chun Ling Zhu ◽  
Su Qing Shi
Keyword(s):  
Heat Exchanger ◽  
Mach Number ◽  
Theoretical Analysis ◽  
Total Pressure ◽  
Experimental Error ◽  
Free Stream ◽  
Test Method ◽  
Flight Altitude ◽  
Simplified Method ◽  
Ground Test

A ground test method of the heat exchanger in intake is proposed using simplified method. A theoretical analysis was carried out to investigate its influences on the performance of heat exchanger. The results show that the performance of heat exchanger keeps invariable with the density of gas if the temperature and mass flow rate remain the same. The error caused by substituting total pressure of free stream for inlet entrance was analyzed and the results show that for subsonic inlet the experimental error is relatively small and the experimental error increasing rapidly as flight Mach number exceed 1.2 and flight altitude has little impact on the experimental error while the error increasing slightly with flight altitude.

Investigation of the Effects of Inlet Swirl on Compressor Performance and Operability Using a Modified Parallel Compressor Model

2011 ◽  
Author(s):  
Nicholas Fredrick ◽  
Milt Davis
Keyword(s):  
Experimental Data ◽  
Total Pressure ◽  
Engine Performance ◽  
Propulsion System ◽  
Turbine Engine ◽  
Turbofan Engines ◽  
Inlet Distortion ◽  
Inlet Swirl ◽  
Engine Compressor ◽  
Compressor Performance

Serpentine ducts used by both military and commercial aircraft can generate significant flow angularity and total pressure distortion at the engine face. Most low by-pass ratio turbofan engines with mixed exhaust are equipped with inlet guide vanes (IGV) which can reduce the effect of moderate inlet distortion. High by-pass ratio and some low by-pass ratio turbofan engines are not equipped with IGVs, and swirl can in effect change the angle of attack of the fan blades. Swirl and total pressure distortion at the engine inlet will impact engine performance, operability, and durability. The impact on the engine performance and operability must be quantified to ensure safe operation of the aircraft and propulsion system. Testing is performed at a limited number of discrete points inside the propulsion system flight envelope where it is believed the engine is most sensitive to the inlet distortion in order to quantify these effects. Turbine engine compressor models are based on the limited amount of experimental data collected during testing. These models can be used as an analysis tool to improve the effectiveness of engine testing and to improve understanding of engine response to inlet distortion. The Dynamic Turbine Engine Compressor Code (DYNTECC) utilizes parallel compressor theory and quasi-one-dimensional Euler equations to determine compressor performance. In its standard form, DYNTECC uses user supplied characteristic stage maps in order to calculate stage forces and shaft work for use in the momentum and energy equations. These maps were typically developed using experimental data or created using characteristic codes such as the 1-D Mean Line Code (MLC) or the 2-D Streamline Curvature Code. The MLC was created to calculate the performance of individual compressor stages and requires less computational effort than the 2-D and 3-D models. To improve efficiency and accuracy, the MLC has been incorporated into DYNTECC as a subroutine. Rather than independently developing stage maps using the MLC and then importing these maps into DYNTECC, DYNTECC can now use the MLC to develop the required stage characteristic for the desired operating point. This will reduce time and complexity required to analyze the effects of inlet swirl on compressor performance. The combined DYNTECC/MLC was used in the past to model total pressure distortion. This paper presents the result obtained using the combined DYNTECC/MLC to model the effects of various types of inlet swirl on F109 fan performance and operability for the first time.

Effect of Non-Uniform Inlet Conditions on Endwall Secondary Flows

2002 ◽  
Author(s):  
K. S. Hermanson ◽  
K. A. Thole
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Total Pressure ◽  
Large Scale ◽  
Secondary Flows ◽  
Turbine Engine ◽  
Thermal Fields ◽  
Pressure Profiles ◽  
Stator Vane ◽  
Inlet Conditions

Exit combustor flow and thermal fields entering downstream stator vane passages in a gas turbine engine are highly non-uniform. These flow and thermal fields can significantly affect the development of the secondary flows in the turbine passages contributing to high platform heat transfer and large aerodynamic losses. The flow and thermal fields combine to give non-uniform total pressure profiles entering the turbine passage which, along with the airfoil geometry, dictate the secondary flow field. This paper presents an analysis of the effects of varying total pressure profiles in both the radial and combined radial and circumferential directions on the secondary flow fields in a first stage stator vane. These inlet conditions used for the first vane simulations are based on the exit conditions predicted for a combustor. Prior to using the predictions, these CFD simulations were benchmarked against flow field data measured in a large-scale, linear, turbine vane cascade. Good agreement occurred between the computational predictions and experimentally measured secondary flows. Analyses of the results for several different cases indicate variations in the secondary flow pattern from pitch to pitch, which attributes to the rationale as to why some airfoils quickly degrade while others remain intact over time.

scholarly journals Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

2012 ◽  
Vol 2012 ◽  
pp. 1-28 ◽  
Author(s):  
Phil Ligrani
Keyword(s):  
Surface Roughness ◽  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
High Speed ◽  
Density Ratio ◽  
Aerodynamic Loss ◽  
Pressure Losses ◽  
Loss Coefficients ◽  
Exit Mach Number

The influences of a variety of different physical phenomena are described as they affect the aerodynamic performance of turbine airfoils in compressible, high-speed flows with either subsonic or transonic Mach number distributions. The presented experimental and numerically predicted results are from a series of investigations which have taken place over the past 32 years. Considered are (i) symmetric airfoils with no film cooling, (ii) symmetric airfoils with film cooling, (iii) cambered vanes with no film cooling, and (iv) cambered vanes with film cooling. When no film cooling is employed on the symmetric airfoils and cambered vanes, experimentally measured and numerically predicted variations of freestream turbulence intensity, surface roughness, exit Mach number, and airfoil camber are considered as they influence local and integrated total pressure losses, deficits of local kinetic energy, Mach number deficits, area-averaged loss coefficients, mass-averaged total pressure loss coefficients, omega loss coefficients, second law loss parameters, and distributions of integrated aerodynamic loss. Similar quantities are measured, and similar parameters are considered when film-cooling is employed on airfoil suction surfaces, along with film cooling density ratio, blowing ratio, Mach number ratio, hole orientation, hole shape, and number of rows of holes.

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.

Designing Forging and Die Tooling for the Manufacture of Turbine Engine Compressor Blades in Siemens NX

2016 ◽  
Vol 684 ◽  
pp. 497-506 ◽  
Author(s):  
D.S. Goryainov ◽  
V.V. Anokhin ◽  
Aleksey Shlyapugin
Keyword(s):  
Point Cloud ◽  
Compressor Blade ◽  
Compressor Blades ◽  
Turbine Engine ◽  
Direct Modeling ◽  
Engine Compressor ◽  
Data Source ◽  
Theoretical Contour ◽  
Gas Turbine Compressor ◽  
Complex Parts

For designing forging and die tooling for bulk forging a necessity in using the data of the geometry of the part produced arises. Obviously, the use as a data source for designing drawings of commonly applied in “manual alternate design” (without a computer) especially such complex parts as compressor blades is not perspective because of the complexity of developing theoretical contour specified by a point cloud. In this case the use of special tooling of direct modeling that provides changing the original model of the part developed by the designers is a perspective one. It should be taken into account during the process of forging and die tooling designing that it is necessary to register the special features of the technology, upon that, the technologist should be highly proficient in using the software. The work given describes the designing technique of gas turbine compressor blade with the account of using the potential of NX Siemens program.

Experiments of Compressible Homogeneous and Isotropic Turbulence Interacting With Expansion Waves

2003 ◽  
Author(s):  
Savvas S. Xanthos ◽  
Yiannis Andreopoulos
Keyword(s):  
Mach Number ◽  
Shock Tube ◽  
Total Pressure ◽  
Large Scale ◽  
Isotropic Turbulence ◽  
Pressure Fluctuations ◽  
Incident Shock ◽  
Expansion Waves ◽  
Curvature Effects ◽  
Homogeneous And Isotropic Turbulence

The interaction of traveling expansion waves with grid-generated turbulence was investigated in a large-scale shock tube research facility. The incident shock and the induced flow behind it passed through a rectangular grid, which generated a nearly homogeneous and nearly isotropic turbulent flow. As the shock wave exited the open end of the shock tube, a system of expansion waves was generated which traveled upstream and interacted with the grid-generated turbulence; a type of interaction free from streamline curvature effects, which cause additional effects on turbulence. In this experiment, wall pressure, total pressure and velocity were measured indicating a clear reduction in fluctuations. The incoming flow at Mach number 0.46 was expanded to a flow with Mach number 0.77 by an applied mean shear of 100 s−1. Although the strength of the generated expansion waves was mild, the effect on damping fluctuations on turbulence was clear. A reduction of in the level of total pressure fluctuations by 20 per cent was detected in the present experiments.

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