Research Papers

Internal Losses and Flow Behavior of a Turbofan Stage at Windmill

[+] Author and Article Information
Dilip Prasad, Wesley K. Lord

Aerodynamics Division, Pratt and Whitney Aircraft Engines, East Hartford, CT 06108

This quantity is defined as the ratio of the capture streamtube area far upstream to the area at the nacelle highlight.

J. Turbomach 132(3), 031007 (Mar 25, 2010) (10 pages) doi:10.1115/1.3147106 History: Received August 26, 2008; Revised September 03, 2008; Published March 25, 2010; Online March 25, 2010

The flow through a high-bypass ratio fan stage during engine-out conditions is investigated, with the objective of quantifying the internal losses when the rotor is at “windmill.” An analysis of altitude test data at various simulated flight Mach numbers shows that the fan rotational speed scales with the engine mass flow rate. Making use of the known values of the nozzle coefficients, we deduce the stagnation pressure loss of the fan stage, which rises significantly as the mass flow rate increases. In order to better understand this behavior, numerical simulations of the fan stage were carried out. The calculated losses agree well with the test data, and it is found that the bulk of the stagnation pressure loss occurs in the stator. A detailed examination of the flow field reveals that the relative flow leaves the rotor at very nearly the metal angle. Moreover, the rotational speed of the fan is such that the inboard sections of the fan blade add work to the flow, while the outboard sections extract work from it. The overall work is essentially zero so that the absolute swirl angle at the rotor exit is small, causing the stator to operate at a severely negative incidence. A gross separation ensues, and the resulting blockage of the stator passage accelerates the flow to high Mach numbers. The highly separated flow in the vane, together with the mixing of the large wakes behind it are responsible for the high losses in the vane. Based on the simulation results for the flow behavior, a simple physical model to estimate the windmill speed of the rotor is developed and is found to be in good agreement with the test data. The utility of this model is that it enables the development of a procedure to predict the internal drag at engine-out conditions, which is discussed.

Copyright © 2010 by American Society of Mechanical Engineers
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Figure 1

(a) Notional meridional section of engine illustrating the stations used in the present study; (b) evolution of the normalized stagnation pressure loss through the fan stage for Φ=0.246

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Figure 2

Variation of dimensionless windmilling tip speed with flight Mach number

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Figure 3

Variation of dimensionless windmilling tip speed with flow parameter. Test data (◻) and model predictions (○, ●) are shown, together with the corresponding lines of second-order best fit (——, – – –). The filled symbols indicate model predictions for points at which test data are not available.

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Figure 4

Stagnation pressure loss as a function of flow parameter. Losses deduced from measurements of the mass flow rate (◇) and thrust (◻) are shown. Simulation results for flow conditions based on the test data (○) and the present model (●) are also indicated. The line represents a second-order fit to the test data, and is extrapolated beyond the original range.

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Figure 5

Stagnation pressure losses across (a) the vane, and (b) the rotor (▽) and splitter (◻) as a function of the flow parameter Φ. The lines represent second-order best fits, while the filled symbols indicate model predictions for points without corresponding test data.

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Figure 6

Velocity triangle and angle convention used in the present analysis. The relative and absolute velocities are denoted by w and c, respectively, and the wheel velocity by U. Flow angles are measured counterclockwise from the positive x-axis.

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Figure 7

Circumferentially-averaged rotor exit radial profiles of (a) relative swirl angle β, and (b) absolute swirl angle α with Φ=0.295 (——), Φ=0.246 (– – –), Φ=0.218(−⋅−⋅−), and Φ=0.194(⋯⋅⋅). Also shown in (a) are the nominal blade exit metal angle β∗ (▽) and the relative flow angle estimated by applying Carter’s rule (◻).

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Figure 8

Circumferentially-averaged radial profiles of the normalized stagnation temperature at the rotor exit with Φ=0.295 (——), Φ=0.246 (– – –), Φ=0.218(−⋅−⋅−), and Φ=0.194(⋯⋅⋅)

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Figure 9

Rotary stagnation pressure at Station RE for Φ=0.246

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Figure 10

Computed blade relative Mach number field at (a) 85% span and (b) 40% span for Φ=0.246

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Figure 11

(a) Numerical (——) and measured (◻) stagnation pressure radial profiles at Station 12.5 for Φ=0.246; (b) numerically determined (——) absolute flow angle and probe angle (◻) as functions of radius

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Figure 12

Computed vane Mach number field at (a) 80% span, (b) 50% span, and (c) 20% span for Φ=0.246

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Figure 13

Normalized stagnation pressure at Station 14 for Φ=0.246. The large blockage caused by the flow separation is evident.

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Figure 14

Vane isentropic Mach number at (a) 80% span, (b) 50% span, and (c) 20% span for Φ=0.246(−⋅−⋅−), Φ=0.295 (——), and Φ=0.313 (– – –)

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Figure 15

Comparison of rotor exit (a) velocity and (b) stagnation temperature profiles obtained using the present model (– – –) with those obtained from the simulations (——) at Φ=0.246

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Figure 16

Illustration of fan stage stagnation pressure loss and nozzle characteristics. The fan operating condition at any given flight Mach number occurs at the intersection of the corresponding nozzle characteristic and the fan loss characteristic.

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Figure 17

Isentropic Mach number distribution (top) over the nacelle surface (bottom) at the windmill conditions listed in Table 1




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