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Research Papers

A Novel Turbomachinery Air-Brake Concept for Quiet Aircraft

[+] Author and Article Information
P. N. Shah

MIT Gas Turbine Laboratory, Massachusetts Institute of Technology, Cambridge, MA 02139parthiv@alum.mit.edu

D. D. Mobed

MIT Gas Turbine Laboratory, Massachusetts Institute of Technology, Cambridge, MA 02139dmobed@alum.mit.edu

Z. S. Spakovszky

MIT Gas Turbine Laboratory, Massachusetts Institute of Technology, Cambridge, MA 02139zolti@mit.edu

This allows for some flow with small but finite radial velocity component to spill or entrain both mass and axial momentum across the outer control volume boundary.

For uniform axial exhaust flows, literature on base flows with bleed (28-29) suggests that the control volume model is valid for velocity ratios above 0.25. Below this velocity ratio vortex shedding occurs.

Note that this is a simple energy constraint only, ignoring the dynamics of vortex breakdown.

Standing wave behavior of the solution indicates this transition (see also Fig. 2).

The circular Burger vortex circulation distribution of Eq. 3 is mapped from radial (r) to streamfunction (ψ) coordinates at the annular actuator disk, under the assumption of uniform axial velocity, i.e., ψ/ψcrit=(r/rcrit)2. This assumption becomes weaker for high levels of swirl, making the effective core radius slightly smaller than the prescribed value. For example, for the swirl tube designs with rcrit specified as 0.7, the maximum value of Vθ occurs near r=0.5. The radial shift of the core is an artifact of the uniform axial velocity assumption, and is the reason that unconverged solutions appear at values of S well below 1.2 for lower values of rc.

This represents an average value of drag coefficients taken at various freestream Mach numbers between M=0.06 and M=0.17.

Taken from a single array microphone.

Directly 105 m overhead, assuming an aircraft on a 3 deg glideslope 2000 m from touchdown.

J. Turbomach 132(4), 041002 (Apr 26, 2010) (11 pages) doi:10.1115/1.3192145 History: Received December 10, 2008; Revised January 16, 2009; Published April 26, 2010; Online April 26, 2010

A novel air-brake concept for next-generation, low-noise civil aircraft is introduced. Deployment of such devices in clean airframe configuration can potentially reduce aircraft source noise and noise propagation to the ground. The generation of swirling outflow from a duct, such as an aircraft engine, is demonstrated to have high drag and low noise. The simplest configuration is a ram pressure-driven duct with stationary swirl vanes, a so-called swirl tube. A detailed aerodynamic design is performed using first principles based modeling and high-fidelity numerical simulations. The swirl-drag-noise relationship is quantified through scale-model aerodynamic and aeroacoustic wind tunnel tests. The maximum measured stable flow drag coefficient is 0.83 at exit swirl angles close to 50 deg. The acoustic signature, extrapolated to full-scale, is found to be well below the background noise of a well-populated area. Vortex breakdown is found to be the aerodynamically and acoustically limiting phenomenon, generating a white-noise signature that is about 15 dB louder than a stable swirling flow.

Copyright © 2010 by American Society of Mechanical Engineers
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References

Figures

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Figure 1

Quiet air-brake concepts: (a) fan-driven or pumped configuration and (b) ram pressure-driven configuration, a so-called swirl tube

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Figure 2

Isocontours of pressure coefficient at super- and subcritical swirl levels. Vortex breakdown instability numerically indicated by waves on vortex core (for illustration purposes the contour density is increased for the swirling exhaust flow).

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Figure 3

Control volume analysis of swirling exhaust flow for drag generation

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Figure 4

Isocontours of drag coefficient and swirl number in Burger vortex design space: (a) control volume analysis, and (b) streamline curvature calculation

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Figure 5

Favorable streamwise pressure gradient along vanes and maximum pressure defect of 3.5 dynamic heads observed in swirl tube with 47 deg vane exit angle setting

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Figure 6

3D CFD calculations: stable exit flow and vortex breakdown, for 47 deg and 57 deg swirl vane angle settings

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Figure 7

Swirl tube mounted in MIT Wright Brothers Wind Tunnel (a) and NASA Langley Quiet Flow Facility (b)

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Figure 8

CFD prediction and measured radial exit profiles of Vx/V∞ and Vθ/V∞ at x/D=1.0 for 34 deg swirl vane angle setting

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Figure 9

CFD and measured swirl parameter profile, S=Vθ/Vx, for 34 deg and 47 degree swirl vane angle settings. Discrepancy seen for 57 deg and 64 deg cases is indicative of vortex breakdown. All data are shown for x/D=1.0.

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Figure 10

Instantaneous flow visualization images show stable swirling flow and vortex breakdown, for swirl vane angle settings of 47 deg and 57 deg, respectively

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Figure 11

Experimentally measured drag coefficient versus swirl angle (model-scale). 3D viscous CFD prediction for converged cases shown for comparison.

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Figure 12

Full-scale noise versus drag versus swirl angle relationship shows high-drag and low-noise capabilities for swirl vane angle setting of 47 deg

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Figure 13

Narrowband (17.44 Hz) spectra for 47 deg versus 57 deg swirl vane angle. Background noise shows good signal-to-noise ratio. Tunnel Mach number is 0.17 in all cases.

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Figure 14

DAMAS zone-integrated spectra and source maps, 47 deg swirl vane angle case, M=0.17, θ=−90 deg. Swirl tube shown in gray in source maps.

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Figure 15

DAMAS zone-integrated spectra and source maps, 57 deg swirl vane angle case, M=0.17, θ=−90 deg. Swirl tube shown in gray in source maps.

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Figure 16

Pumped swirl tube configuration in thrust reverser blocker door mode

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