The Effect of Heat Transfer on Gas Turbine Transients

This paper describes how allowance for the thermal effects of non-adiabatic flow, altered boundary layer development, changes in tip clearances and changes in seal clearances have been incorporated into a general gas turbine transient program. These non-adiabatic effects have been investigated, modelling a two-spool bypass engine. The model has predicted events that occur in practice and also indicates which of the parameters are the most influential in the alteration of transient performance.

Volume 1: Aircraft Engine; Marine; Turbomachinery; Microturbines and Small Turbomachinery ◽

The Prediction of Surge Margins During Gas Turbine Transients

10.1115/85-gt-208 ◽

1985 ◽

Cited By ~ 1

Author(s):

N. R. L. Maccallum ◽

P. Pilidis

Keyword(s):

Boundary Layer ◽

Gas Turbine ◽

Thermal Effects ◽

Adiabatic Flow ◽

Surge Margin ◽

Surge Line ◽

This paper describes how allowance for the thermal effects of non-adiabatic flow, altered boundary layer development, changes in tip clearances and changes in seal clearances have been incorporated into a general gas turbine transient program. This program has been applied to a two-spool bypass engine. Revised predictions of surge margins in three common transients have been obtained. When the engine undergoes a “cold” acceleration, the thermal effects on the trajectory and on the surge line give a much increased proportion of unused surge margin in the H.P. Compressor, as compared to adiabatic predictions. In a “hot” acceleration this improvement is considerably reduced.

Numerical Simulation of Turbine Blade Boundary Layer and Heat Transfer and Assessment of Turbulence Models

Journal of Turbomachinery ◽

10.1115/1.2841190 ◽

1997 ◽

Vol 119 (4) ◽

pp. 794-801 ◽

Cited By ~ 7

Author(s):

J. Luo ◽

B. Lakshminarayana

Keyword(s):

Heat Transfer ◽

Boundary Layer ◽

Reynolds Number ◽

Low Reynolds Number ◽

Turbulence Models ◽

Navier Stokes ◽

Bypass Transition ◽

Low Reynolds ◽

The boundary layer development and convective heat transfer on transonic turbine nozzle vanes are investigated using a compressible Navier–Stokes code with three low-Reynolds-number k–ε models. The mean-flow and turbulence transport equations are integrated by a four-stage Runge–Kutta scheme. Numerical predictions are compared with the experimental data acquired at Allison Engine Company. An assessment of the performance of various turbulence models is carried out. The two modes of transition, bypass transition and separation-induced transition, are studied comparatively. Effects of blade surface pressure gradients, free-stream turbulence level, and Reynolds number on the blade boundary layer development, particularly transition onset, are examined. Predictions from a parabolic boundary layer code are included for comparison with those from the elliptic Navier–Stokes code. The present study indicates that the turbine external heat transfer, under real engine conditions, can be predicted well by the Navier–Stokes procedure with the low-Reynolds-number k–ε models employed.

Some Boundary Layer-Porous Wall Coupling Effects in Laminar Flows With Surface Injection

Journal of Engineering for Power ◽

10.1115/1.3445350 ◽

1970 ◽

Vol 92 (3) ◽

pp. 257-266

Author(s):

D. A. Nealy ◽

P. W. McFadden

Keyword(s):

Heat Transfer ◽

Boundary Layer ◽

Heat Balance ◽

Transfer Problem ◽

Porous Wall ◽

Laminar Flows ◽

Stream Velocity ◽

Axial Conduction ◽

Using the integral form of the laminar boundary layer thermal energy equation, a method is developed which permits calculation of thermal boundary layer development under more general conditions than heretofore treated in the literature. The local Stanton number is expressed in terms of the thermal convection thickness which reflects the cumulative effects of variable free stream velocity, surface temperature, and injection rate on boundary layer development. The boundary layer calculation is combined with the wall heat transfer problem through a coolant heat balance which includes the effect of axial conduction in the wall. The highly coupled boundary layer and wall heat balance equations are solved simultaneously using relatively straightforward numerical integration techniques. Calculated results exhibit good agreement with existing analytical and experimental results. The present results indicate that nonisothermal wall and axial conduction effects significantly affect local heat transfer rates.

Boundary Layer Development under Pressure Gradients, with Particular Reference to Heat Transfer

Theory and Fundamental Research in Heat Transfer ◽

10.1016/b978-0-08-009936-1.50013-9 ◽

1963 ◽

pp. 161-180

Author(s):

R.J. MONAGHAN

Keyword(s):

Heat Transfer ◽

Boundary Layer ◽

Pressure Gradients ◽

Investigation into convective heat transfer of nanofluids at the thermal boundary layer development region in micro-channels

IOP Conference Series Materials Science and Engineering ◽

10.1088/1757-899x/793/1/012043 ◽

2020 ◽

Vol 793 ◽

pp. 012043

Author(s):

Shuxiang Wang ◽

Li Xu ◽

Junjie Tong

Keyword(s):

Heat Transfer ◽

Boundary Layer ◽

Convective Heat Transfer ◽

Convective Heat ◽

Thermal Boundary Layer ◽

Micro Channels ◽

Development Region ◽

Free-Stream Turbulence and Pressure Gradient Effects on Heat Transfer and Boundary Layer Development on Highly Cooled Surfaces

Journal of Engineering for Gas Turbines and Power ◽

10.1115/1.3239697 ◽

1985 ◽

Vol 107 (1) ◽

pp. 54-59 ◽

Cited By ~ 6

Author(s):

K. Rued ◽

S. Wittig

Keyword(s):

Heat Transfer ◽

Boundary Layer ◽

Free Stream ◽

Boundary Layer Transition ◽

Pressure Gradients ◽

Layer Transition ◽

Free Stream Turbulence ◽

Gradient Effects ◽

Heat transfer and boundary layer measurements were derived from flows over a cooled flat plate with various free-stream turbulence intensities (Tu = 1.6–11 percent), favorable pressure gradients (k = νe/ue2•due/dx = 0÷6•10−6) and cooling intensities (Tw/Te = 1.0–0.53). Special interest is directed towards the effects of the dominant parameters, including the influence on laminar to turbulent boundary layer transition. It is shown, that free-stream turbulence and pressure gradients are of primary importance. The increase of heat transfer due to wall cooling can be explained primarily by property variations as transition, and the influence of free-stream parameters are not affected.

HEAT TRANSFER CHARACTERISTICS AND BOUNDARY LAYER DEVELOPMENT ABOUT HEATING AND COOLING ROTATING BLUNT BODIES AT SUPERSONIC SPEEDS

Proceeding of International Heat Transfer Conference 7 ◽

10.1615/ihtc7.1610 ◽

1982 ◽

Author(s):

David C. Chou ◽

James A. Kalina

Keyword(s):

Heat Transfer ◽

Boundary Layer ◽

Heat Transfer Characteristics ◽

Transfer Characteristics ◽

Heating And Cooling ◽

Blunt Bodies ◽

Modeling Thermal Effects on Performance of Small Gas Turbines

Volume 1: Aircraft Engine; Fans and Blowers; Marine ◽

10.1115/gt2015-42744 ◽

2015 ◽

Cited By ~ 1

Author(s):

W. P. J. Visser ◽

I. D. Dountchev

Keyword(s):

Heat Transfer ◽

Steady State ◽

Gas Turbine ◽

Thermal Effects ◽

Gas Turbines ◽

Engine Performance ◽

Unmanned Aircraft ◽

Transient Performance ◽

Large Engine ◽

Aircraft Propulsion

Gas turbines are applied at increasingly smaller scales for both aircraft propulsion and power generation. Recuperated turboshaft micro turbines below 30 kW are being developed at efficiencies competitive with other heat engines. The rapidly increasing number of unmanned aircraft applications requires the development of small efficient aircraft propulsion gas turbines. Thermal effects such as steady-state heat losses and transient heat soakage on large engine performance are relatively small and therefore often neglected in performance simulations. At small scales however, these become very significant due to the much higher heat transfer area-to-volume ratios in the gas path components. Recuperators often have high heat capacity and therefore affect transient performance significantly, also with large engine scales. As a result, for accurate steady-state and transient performance prediction of micro and recuperated gas turbines, thermal effects need to be included with sufficient fidelity. In the paper, a thermal network model functionality is presented that can be integrated in a gas turbine system simulation environment such as the Gas turbine Simulation Program GSP [1]. In addition, a 1-dimensional thermal effects model for recuperators is described. With these two elements, thermal effects in small recuperated gas turbines can be accurately predicted. Application examples are added demonstrating and validating the methods with models of a recuperated micro turbine. Simulation results are given predicting effects of heat transfer and heat loss on steady-state and transient performance.

An Algebraic Model for High Intensity Large Scale Turbulence

Volume 3: Heat Transfer; Electric Power; Industrial and Cogeneration ◽

10.1115/99-gt-160 ◽

1999 ◽

Cited By ~ 4

Author(s):

F. E. Ames ◽

O. Kwon ◽

R. J. Moffat

Keyword(s):

Heat Transfer ◽

Boundary Layer ◽

Turbulent Boundary Layer ◽

Outer Layer ◽

High Intensity ◽

Large Scale ◽

Layer Model ◽

External Turbulence ◽

An algebraic turbulence model, which has been developed based on the dynamics of ν′ spectra of external turbulence near a surface, is presented in this paper. The model provides an accurate method of predicting the influence of large-scale high intensity turbulence on heat transfer and boundary layer development in turbomachinery. The model has been developed to predict both laminar and turbulent boundary layer development and heat transfer. The laminar boundary layer model has been tested against boundary layer data taken in a low speed cascade. The model produces accurate velocity and eddy diffusivity distributions. The turbulent boundary layer model is composed of inner and outer layer models combined with an intermittency function. The inner model is written in the form of a conventional mixing length model; while the outer layer model is expressed in terms of the external turbulence characteristics. Predictions of boundary layer profiles and heat transfer distributions are shown for both turbulent and laminar boundary layers. Vane Stanton number predictions were made for inlet turbulence levels ranging from one to thirteen percent for a chord Reynolds number of 800,000. Predictions agreed with experimentally determined levels within 6 percent on the pressure surface but were underpredicted by up to 15 percent in the stagnation region. Levels of heat transfer predicted in the turbulent region of the suction surface agreed with the data within 10 percent.