Gas Turbine Laboratory

Montreal, United States

Gas Turbine Laboratory

Montreal, United States

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Currie T.C.,National Research Council Canada | Currie T.C.,Gas Turbine Laboratory | Struk P.M.,NASA | Tsao J.-C.,Ohio Aerospace Institute | And 4 more authors.
4th AIAA Atmospheric and Space Environments Conference 2012 | Year: 2012

This paper describes experiments performed in an altitude chamber at the National Research Council of Canada (NRC) as the first phase of a joint NRC/NASA program investigating ice crystal accretion in aero engines. The principal objective was to explore the effect of wet bulb temperature Twb (dependent on air temperature, humidity and pressure) on accretion behavior, since preliminary results published in an earlier paper indicated that well-adhered accretions are only possible at Twb<0°C, when water in an impinging mixedphase flow can freeze to a surface. To assess the accretion sensitivity to Twb, the symmetrical airfoil used in the previous work was tested at pressures of 44.8 kPa and 93kPa, usually at 0.25 Mach number, over a range of freestream liquid water and ice water concentrations, total air temperatures and humidity levels. Twb was typically maintained at +2°C or -2°C, based on dry total conditions (i.e. without ice or water injection). Total air temperature was >0°C in all tests. The limited test results confirmed that accretion behavior is very sensitive to Twb, which is in turn strongly related to pressure since evaporative cooling increases with decreasing pressure. Humidity and total temperature did not appear to have an independent effect on accretion behavior. Accretions, often resembling glaze ice, formed at Twb<0°C, when freestream water would freeze on the test airfoil without ice crystals present in the freestream. At Twb>0°C ice deposits were observed to be slushy, poorly adhered and shed frequently. The size of such deposits appeared to be a non-linear function of the freestream ice water content (IWC), becoming much larger at high IWC. © 2012 by National Research Council of Canada. Published by the American Institute of Aeronautics and Astronautics, Inc.

Knezevici D.C.,National Research Council Canada | Knezevici D.C.,Gas Turbine Laboratory | Fuleki D.,National Research Council Canada | Fuleki D.,Gas Turbine Laboratory | And 4 more authors.
4th AIAA Atmospheric and Space Environments Conference 2012 | Year: 2012

This paper describes the commissioning of a new test apparatus intended to simulate an inner-compressor duct bleed slot. It also identifies, for the first time, that ice crystal particle size plays an important role in the ice crystal phenomenon. Data and sample images of accretion are presented for wet bulb temperatures near freezing. The effect of wet bulb temperature and particle size on the natural melting of ice crystals is investigated. In addition, the erosion of surface accretion by ice crystal particles is discussed. © 2012 by Her Majesty the Queen in Right of Canada. Published by the American Institute of Aeronautics and Astronautics, Inc.

Currie T.C.,National Research Council Canada | Currie T.C.,Principal Research Officer | Fuleki D.,National Research Council Canada | Fuleki D.,Gas Turbine Laboratory
8th AIAA Atmospheric and Space Environments Conference | Year: 2016

Aircraft jet engines operating at high altitudes in ice crystal clouds can experience operational problems and/or damage resulting from accretion of ingested ice crystals within the compressor. It is believed the ice crystals partially melt, allowing them to stick to internal components. Most previous research into such ice crystal icing (ICI) conducted at the National Research Council of Canada (NRCC) has used low Mach numbers (M), usually <0.4, whereas the axial Mach number within an LP compressor is typically 0.5-0.6, and blade surface Mach numbers are even higher. It has been hypothesized that accretion at such high M probably requires particles with an MVD <30μ, which is smaller than used previously. This paper describes high M ICI tests conducted with hemispherical and double wedge geometries in the NRCC RATFac icing tunnel to test this hypothesis. These tests used 28μ MVD ice particles produced with a grinder and relied on natural melting of the particles in warm air. In contrast to earlier tests of these geometries conducted with larger (57μ MVD) particles in which no accretion was observed above M=0.4, accretion was observed with the smaller particles at M up to 0.59 (wedge) and 0.65 (hemisphere). Accretion in such cases initiated at a location on the geometry where the local wet bulb temperature based on the local recovery temperature and static pressure dropped below freezing, as predicted by CFD. Accretions were conical, as in earlier tests. The paper also describes ICI tests performed with the hemispherical geometry at sub-freezing temperatures, using liquid water injected with spray nozzles. These tests used 57μ MVD particles and M=0.25 to permit comparison with earlier tests which relied on natural melt at warmer temperatures and otherwise the same conditions. Whereas the latter produced poorly adhered conical accretions which grew forwards with a gradually decreasing cone angle and a base which remained tangent to the hemisphere, the sub-freezing tests produced well-adhered conical accretions with a constant cone angle (typically) and a flared mushroom-shaped base. The sticking efficiency model published previously has been modified to predict the high M/small particle and sub-freezing Twbo accretions with reasonable fidelity. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA.

Ng L.W.T.,Massachusetts Institute of Technology | Ng L.W.T.,Gas Turbine Laboratory | Spakovszky Z.,Massachusetts Institute of Technology | Spakovszky Z.,Gas Turbine Laboratory
AIAA Journal | Year: 2011

This paper presents a method based on the Kirchhoff diffraction theory to predict the shielding of turbomachinery noise by the airframe of an advanced aircraft configuration. The key feature of this method is the fast computational time, even at very high frequencies, which makes it a useful tool to rapidly assess the noise footprint of an aircraft design. The offline part preprocesses the three-dimensional shielding geometry into a contour of its outline based on the source line of sight. Given this contour, the online part calculates the noise attenuation at a particular observer location and source frequency and can be called multiple times by an aircraft noise prediction program to add shielding estimates to its effective perceived noise level calculations. This method is most accurate for flat shielding objects characterized by edge-diffraction rays, rather than smooth, rounded objects characterized by creeping rays; shielding differences of up to 3 dB were observed in calculations using a sphere and a disk. Finally, the method is applied to a hybrid wing-body aircraft to assess and quantify the noise shielding benefit of its large planform area. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Hileman J.I.,Massachusetts Institute of Technology | Hileman J.I.,Gas Turbine Laboratory | Spakovszky Z.S.,Massachusetts Institute of Technology | Spakovszky Z.S.,Gas Turbine Laboratory | And 4 more authors.
Journal of Aircraft | Year: 2010

The noise goal of the Silent Aircraft Initiative, a collaborative effort between the University of Cambridge and Massachusetts Institute of Technology, demanded an airframe design with noise as a prime design variable and a design philosophy that cut across multiple disciplines. This paper discusses a novel design methodology synthesizing first-principles analysis and high-fidelity simulations, and it presents the conceptual design of an aircraft with a calculated noise level of 62 dBA at the airport perimeter. This is near the background noise in a well-populated area, making the aircraft imperceptible to the human ear on takeoff and landing. The all-lifting airframe of the conceptual aircraft design also has the potential for improved fuel efficiency, as compared with existing commercial aircraft. A key enabling technology in this conceptual design is the aerodynamic shaping of the airframe centerbody. Design requirements and challenges are identified, and the resulting aerodynamic design is discussed in depth. The paper concludes with suggestions for continued research on enabling technologies for quiet commercial aircraft. Copyright © 2010 by the authors.

Shah P.N.,Massachusetts Institute of Technology | Shah P.N.,Gas Turbine Laboratory | Mobed D.,Massachusetts Institute of Technology | Mobed D.,Gas Turbine Laboratory | And 4 more authors.
AIAA Journal | Year: 2010

Aircraft on approach in high-drag, high-lift configurations create inherently noisy flow structures. For flaps, slats, and undercarriage, the strong correlation between overall noise and drag suggests that future quiet aircraft will need to generate drag at low noise levels. This paper presents a novel noise-reduction concept based on the idea that appreciable pressure drag can be generated by a relatively quiet swirling exhaust flow. A first aeroacoustic assessment of ram-pressure-driven swirling exhaust flows and their associated vortex breakdown instability is presented. The technical approach combines 1) an in-depth aerodynamic analysis, 2) qualitative acoustic source descriptions via plausibility arguments, and 3) detailed quantitative phased microphone-array measurements of a model-scale engine nacelle with stationary swirl vanes at a full-scale approach Mach number of 0.17. The analysis shows an acoustic signature composed of 1) quadrupole-type turbulent mixing noise in the swirling core flow and 2) scattering noise from vane boundary layers and turbulent eddies of the burst vortex structure near the nacelle, pylon, and vane centerbody trailing edges. The highest stable swirl-angle setting yields a nacelle-area-based drag coefficient of 0.83 with a full-scale overall sound pressure level of about 40 dBA at the International Civil Aviation Organization approach certification point. Copyright © 2009 by P. N. Shah, D. Mobed, Z. S. Spakovszky, T. F. Brooks, and W. M. Humphreys Jr.

Tiralap A.,Massachusetts Institute of Technology | Tan C.S.,Massachusetts Institute of Technology | Tan C.S.,Gas Turbine Laboratory | Donahoo E.,Siemens AG | And 2 more authors.
Proceedings of the ASME Turbo Expo | Year: 2016

Changes in loss generation associated with altering the rotor tip blade loading of an embedded rotor-stator compressor stage are assessed with unsteady three-dimensional computations, complemented by control volume analyses. Tip-foreloaded and tip-aft-loaded rotor blades are designed and assessed to provide variation in rotor tip blade loading distributions for determining if aft-loading rotor tip would yield a stage performance benefit in terms of a reduction in loss generation. Aftloading rotor blade tip delays the formation of tip leakage flow resulting in a relatively less mixed-out tip leakage flow at the rotor outlet and a reduction in overall tip leakage mass flow, hence a lower loss generation; however, the attendant changes in tip flow angle distribution are such that there is an overall increase in the flow angle mismatch between tip flow and main flow leading to higher loss generation. The latter outweighs the former so that rotor passage loss from aft-loading rotor tip is marginally higher unless a constraint is imposed on tip flow angle distribution so that associated induced loss is negligible; a potential strategy for achieving this is proposed. Tip leakage flow, which is not mixed-out at the rotor outlet, enters the downstream stator, where it can be recovered. The tip leakage flow recovery process yields a higher benefit for a relatively less mixed-out tip leakage flow from aft-loading a rotor blade tip. These characterizing parameters together determine the attendant loss associated with rotor tip leakage flow in a compressor stage environment. A revised design hypothesis is thus as follows: Rotor should be tip-aft-loaded and hub-fore-loaded while stator should be hubaft- loaded and tip-fore-loaded with tip/hub leakage flow angle distribution such that it results in no additional loss. For the compressor stage being assessed here, an estimated 0.15 points enhancement in stage efficiency is possible from aft-loading rotor tip only. In the course of assessing the benefit from unsteady tip leakage flow recovery in the downstream stator, it was determined that tip clearance flow is inherently unsteady with a time-scale distinctly different from the blade passing time. The disparity between the two timescales: (i) defines the periodicity of the unsteady rotor-stator flow, which is an integral multiple of blade passing time; and (ii) causes tip leakage vortex to enter the downstream stator at specific pitchwise locations for different blade passing cycles, a tip leakage flow phasing effect. Despite the inherent unsteadiness from tip leakage flow, the recovery process is demonstrated to be beneficial on a time-averaged basis. © 2016 by ASME.

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