Experimental Reduction of Transonic Fan Forced Response by Inlet Guide Vane Flow Control

2009 ◽  
Vol 132 (2) ◽  
Author(s):  
S. Todd Bailie ◽  
Wing F. Ng ◽  
William W. Copenhaver
Keyword(s):  
Flow Control ◽  
Guide Vane ◽  
Leading Edge ◽  
Forced Response ◽  
Inlet Guide Vane ◽  
Resonant Response ◽  
Compressor Blades ◽  
Resonant Amplitude ◽  
Blade Vibrations ◽  
Transonic Fan

The main contributor to the high cycle fatigue of compressor blades is the response to aerodynamic forcing functions generated by an upstream row of stators or inlet guide vanes. Resonant response to engine order excitation at certain rotor speeds can be especially damaging. Studies have shown that flow control by trailing edge blowing (TEB) can reduce stator wake strength and the amplitude of the downstream rotor blade vibrations generated by the unsteady stator-rotor interaction. In the present study, the effectiveness of TEB to reduce forced fan blade vibrations was evaluated in a modern single-stage transonic fan rig. Data were collected for multiple uniform full-span TEB conditions over a range of rotor speeds including multiple modal resonance crossings. Resonant response sensitivity was generally characterized by a robust region of strong attenuation. The baseline resonant amplitude of the first torsion mode, which exceeded the endurance limit on the critical blade, was reduced by more than 80% with TEB at 1.0% of the total rig flow. The technique was also found to be modally robust; similar reductions were achieved for all tested modal crossings, including more than 90% reduction in the second leading-edge bending response using 0.7% of the rig flow.

Experimental Reduction of Transonic Fan Forced Response by IGV Flow Control

2004 ◽  
Author(s):  
S. Todd Bailie ◽  
Wing F. Ng ◽  
William W. Copenhaver
Keyword(s):  
Flow Control ◽  
Forced Response ◽  
Resonant Response ◽  
Compressor Blades ◽  
Resonant Amplitude ◽  
Engine Order ◽  
Blade Vibrations ◽  
Fan Blade ◽  
Transonic Fan ◽  
Bending Response

The main contributor to the high-cycle fatigue of compressor blades is the response to aerodynamic forcing functions generated by an upstream row of stators or inlet guide vanes. Resonant response to engine order excitation at certain rotor speeds can be especially damaging. Studies have shown that flow control by trailing edge blowing (TEB) can reduce stator wake strength and the amplitude of the downstream rotor blade vibrations generated by the unsteady stator-rotor interaction. In the present study, the effectiveness of TEB to reduce forced fan blade vibrations was evaluated in a modern single-stage transonic fan rig. Data was collected for multiple uniform full-span TEB conditions over a range of rotor speed including multiple modal resonance crossings. Resonant response sensitivity was generally characterized by a robust region of strong attenuation. The baseline resonant amplitude of the first torsion mode, which exceeded the endurance limit on the critical blade, was reduced by more than 80% with TEB at 1.0% of the total rig flow. The technique was also found to be modally robust; similar reductions were achieved for all tested modal crossings, including more than 90% reduction of the second LE bending response using 0.7% of the rig flow.

Influence of Gap Detailing on Calculated Unsteady Non-Adjacent Blade Row Aero-Forcing in a Transonic Compressor Stage

2020 ◽  
Vol 142 (11) ◽  
Author(s):  
Tobias Gezork ◽  
Paul Petrie-Repar
Keyword(s):  
Guide Vane ◽  
Computational Cost ◽  
Life Time ◽  
Forced Response ◽  
Inlet Guide Vane ◽  
Compressor Blades ◽  
Transonic Compressor ◽  
Forcing Function ◽  
Gap Flow ◽  
Blade Row

Abstract Resonant or close to resonant forced response excitation of compressor blades limits component life time and can potentially lead to high-cycle fatigue failure if the exciting forces are large and damping is insufficient. When numerically quantifying the forcing function by means of simulations, simplifications are typically made in the analysis to reduce complexity and computational cost. In this paper, we numerically investigate how the blade forcing function is influenced by the rotor tip gap flow and by flow across gaps in the upstream variable inlet guide vane row. Unsteady simulations are made using a test rig geometry where a forcing crossing with an excitation from a non-adjacent blade row had previously been measured. The effects of the gaps on the forcing function for the first torsion mode are presented for both the non-adjacent blade row excitation (changes compared with a case without gaps indicating a 20% reduction) and an adjacent excitation (changes indicating an 80% increase in terms of forcing function amplitude comparing with a case without gaps).

Effects of Flow Control on Forced Response and Performance of a Transonic Compressor

2002 ◽  
Author(s):  
S. Todd Bailie ◽  
Wing F. Ng ◽  
Alfred L. Wicks ◽  
William W. Copenhaver
Keyword(s):  
Flow Control ◽  
Rotor Blade ◽  
Forced Response ◽  
Strain Gages ◽  
Complex Flow ◽  
Compressor Blades ◽  
Transonic Compressor ◽  
Engine Order ◽  
Blade Vibrations ◽  
And Performance

The main contributor to the high-cycle fatigue of compressor blades is the response to aerodynamic forcing functions generated by an upstream row of stators or inlet guide vanes. Resonant response to engine order excitation at certain rotor speeds is especially damaging. Studies have shown that flow control by trailing edge blowing (TEB) can reduce stator wake strength and the amplitude of the downstream rotor blade vibrations generated by the unsteady stator-rotor interaction. In the present study, the effectiveness of TEB to reduce forced blade vibrations was evaluated in a modern transonic compressor rig. A row of wake generator (WG) vanes with TEB capability was installed upstream of the rotor, which was instrumented with strain gages. Data was collected with and without TEB at various rotor speeds involving resonance crossings. Using 0.8% of the compressor core flow for TEB along the full WG-span, rotor blade strain was reduced by 66% at the first torsional resonance crossing. Substantial reductions were also achieved with only partial span TEB. The results demonstrate the effectiveness of the TEB technique for reducing rotor vibrations in the complex flow environment of a closely-spaced transonic stage row. Moderate increases in stage performance were also measured.

scholarly journals Redesigning a compressor inlet guide vane for 0 to 60 degree stagger angles

2021 ◽  
Author(s):  
Stephanie Waters
Keyword(s):  
Fuel Consumption ◽  
Guide Vane ◽  
Leading Edge ◽  
Aircraft Engine ◽  
Inlet Guide Vane ◽  
Pressure Loss Coefficient ◽  
Compressor Inlet ◽  
Curvature Distribution ◽  
Stagger Angle ◽  
Total Pressure Loss Coefficient

This report's objective is to reduce the total pressure loss coefficient of an inlet guide vane (IGV) at high stagger angles and to therefore reduce the overall fuel consumption of an aircraft engine. IGVs are usually optimized for cruise where the stagger angle is approximately 0 degrees. To reduce losses, four different methodologies were tested: increasing the leading edge radius, increasing the camber, creating a "drooped nose", and creating an "S" curvature distribution. A baseline IGV was chosen and modified using these methodologies to create 10 new IGV designs. CFX was used to perform a CFD analysis on all 11 IGV designs at 5 stagger angles from 0 to 60 degrees. Typical missions were analyzed and it was discovered that the new designs decreased the fuel consumption of the engine. The IGV with the "S" curvature and thicker leading edge was the best and decreased the fuel consumption by 0.24%.

Ice Crystal Icing Investigation on a Honeywell Uncertified Research Engine in an Altitude Simulation Icing Facility

2020 ◽  
Author(s):  
Ashlie B. Flegel
Keyword(s):  
Particle Size ◽  
Guide Vane ◽  
Leading Edge ◽  
Inlet Guide Vane ◽  
Inlet Temperature ◽  
Ice Crystal ◽  
Guide Vanes ◽  
Inlet Guide Vanes ◽  
The Core ◽  
Break Up

Abstract A Honeywell Uncertified Research Engine was exposed to various ice crystal conditions in the NASA Glenn Propulsion Systems Laboratory. Simulations using NASA’s 1D Icing Risk Analysis tool were used to determine potential inlet conditions that could lead to ice crystal accretion along the inlet of the core flowpath and into the high pressure compressor. These conditions were simulated in the facility to develop baseline conditions. Parameters were then varied to move or change accretion characteristics. Data were acquired at altitudes varying from 5 kft to 45 kft, at nominal ice particle Median Volumetric Diameters from 20 μm to 100 μm, and total water contents of 1 g/m3 to 12 g/m3. Engine and flight parameters such as fan speed, Mach number, and inlet temperature were also varied. The engine was instrumented with total temperature and pressure probes. Static pressure taps were installed at the leading edge of the fan stator, front frame hub, the shroud of the inlet guide vane, and first two rotors. Metal temperatures were acquired for the inlet guide vane and vane stators 1–2. In-situ measurements of the particle size distribution were acquired three meters upstream of the engine forward fan flange and one meter downstream of the fan in the bypass in order to study particle break-up behavior. Cameras were installed in the engine to capture ice accretions at the leading edge of the fan stator, splitter lip, and inlet guide vane. Additional measurements acquired but not discussed in this paper include: high speed pressure transducers installed at the trailing edge of the first stage rotor and light extinction probes used to acquire particle concentrations at the fan exit stator plane and at the inlet to the core and bypass. The goal of this study was to understand the key parameters of accretion, acquire particle break-up data aft of the fan, and generate a unique icing dataset for model and tool development. The work described in this paper focuses on the effect of particle break-up. It was found that there was significant particle break-up downstream of the fan in the bypass, especially with larger initial particle sizes. The metal temperatures on the inlet guide vanes and stators show a temperature increase with increasing particle size. Accretion behavior observed was very similar at the fan stator and splitter lip across all test cases. However at the inlet guide vanes, the accretion decreased with increasing particle size.

scholarly journals Investigation on the Flow Field Entropy Structure of Non-Synchronous Blade Vibration in an Axial Turbocompressor

Entropy ◽  
2020 ◽  
Vol 22 (12) ◽  
pp. 1372
Author(s):  
Mingming Zhang ◽  
Anping Hou
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Guide Vane ◽  
Leading Edge ◽  
Inlet Guide Vane ◽  
Tip Leakage Flow ◽  
Good Correspondence ◽  
Blade Vibration ◽  
Tip Leakage ◽  
Spiral Vortex

In order to explore the inducing factors and mechanism of the non-synchronous vibration, the flow field structure and its formation mechanism in the non-synchronous vibration state of a high speed turbocompressor are discussed in this paper, based on the fluid–structure interaction method. The predicted frequencies fBV (4.4EO), fAR (9.6EO) in the field have a good correspondence with the experimental data, which verify the reliability and accuracy of the numerical method. The results indicate that, under a deviation in the adjustment of inlet guide vane (IGV), the disturbances of pressure in the tip diffuse upstream and downstream, and maintain the corresponding relationship with the non-synchronous vibration frequency of the blade. An instability flow that developed at the tip region of 90% span emerged due to interactions among the incoming main flow, the axial separation backflow, and the tip leakage vortices. The separation vortices in the blade passage mixed up with the tip leakage flow reverse at the trailing edge of blade tip, presenting a spiral vortex structure which flows upstream to the leading edge of the adjacent blade. The disturbances of the spiral vortexes emerge to rotate at 54.5% of the rotor speed in the same rotating direction as a modal oscillation. The blade vibration in the turbocompressor is found to be related to the unsteadiness of the tip flow. The large pressure oscillation caused by the movement of the spiral vortex is regarded as the one of the main drivers for the non-synchronous vibration for the present turbocompressor, besides the deviation in the adjustment of IGV.

Flow Control on a Transport Truck Side Mirror Using Plasma Actuators

2015 ◽  
Vol 137 (11) ◽  
Author(s):  
Theodoros Michelis ◽  
Marios Kotsonis
Keyword(s):  
Flow Control ◽  
Guide Vane ◽  
Leading Edge ◽  
Plasma Actuator ◽  
Plasma Actuators ◽  
Dbd Plasma ◽  
Reference Case ◽  
Image Velocimetry ◽  
Wind Tunnel Study ◽  
Edge Separation

A wind tunnel study is conducted toward hybrid flow control of a full scale transport truck side mirror at ReD=3.2×105. A slim guide vane is employed for redirecting high-momentum flow toward the mirror wake region. Leading edge separation from the guide vane is reduced or eliminated by means of an alternating current -dielectric barrier discharge (AC-DBD) plasma actuator. Particle image velocimetry (PIV) measurements are performed at a range of velocities from 15 to 25 m/s and from windward to leeward angles from -5deg to 5deg. Time-averaged velocity fields are obtained at the center of the mirror for three scenarios: (a) reference case lacking any control elements, (b) guide vane only, and (c) combination of the guide vane and the AC-DBD plasma actuator. The comparison of cases demonstrates that at 25 m/s windward conditions (-5deg) the guide vane is capable of recovering 17% momentum with respect to the reference case. No significant change is observed by activating the AC-DBD plasma actuator. In contrast, at leeward conditions (5deg), the guide vane results in a −20% momentum loss that is rectified to a 6% recovery with actuation. The above implies that for a truck with two mirrors, 23% of momentum may be recovered.

Periodic Transition on an Axial Compressor Stator — Incidence and Clocking Effects: Part I — Experimental Data

1998 ◽  
Author(s):  
G. J. Walker ◽  
J. D. Hughes ◽  
W. J. Solomon
Keyword(s):  
Guide Vane ◽  
Axial Compressor ◽  
Separated Flow ◽  
Leading Edge ◽  
Inlet Guide Vane ◽  
Turbulent Spot ◽  
Pressure Surface ◽  
Surface Pressure Distribution ◽  
Surface Transition ◽  
Hot Film

Periodic wake-induced transition on the outlet stator of a 1.5 stage axial compressor is examined using hot-film arrays on both the suction and pressure surfaces. The time-mean surface pressure distribution is varied by changing the blade incidence, while the freestream disturbance field is altered by clocking of the stator relative to an inlet guide vane row. Ensemble average plots of turbulent intermittency and relaxation factor (extent of calmed flow following the passage of a turbulent spot) are presented. These show the strength of periodic wake-induced transition phenomena to be significantly influenced by both incidence and clocking effects. The nature and extent of transition by other modes (natural, bypass and separated flow transition) are altered accordingly. Leading edge and mid-chord separation bubbles are affected in a characteristically different manner by changing freestream periodicity. There are noticeable differences between suction and pressure surface transition behavior, particularly as regards the strength and extent of calming. In Part II of this paper, the transition onset observations from the compressor stator are used to evaluate the quasi-steady application of conventional transition correlations to predict unsteady transition onset on the blading of an embedded axial compressor stage.

ICE CRYSTAL ICING INVESTIGATION ON A HONEYWELL UNCERTIFIED RESEARCH ENGINE IN AN ALTITITUDE SIMULATION ICING FACILITY

2021 ◽  
pp. 1-22
Author(s):  
Ashlie Flegel
Keyword(s):  
Particle Size ◽  
Guide Vane ◽  
Model Development ◽  
Leading Edge ◽  
Inlet Guide Vane ◽  
Analysis Tool ◽  
Ice Crystal ◽  
High Pressure Compressor ◽  
Break Up ◽  
Inlet Conditions

Abstract A Honeywell Uncertified Research Engine was exposed to various ice crystal conditions in the NASA Propulsion Systems Laboratory. Simulations using NASA's 1D Icing Risk Analysis tool were used to determine potential inlet conditions that could lead to ice crystal accretion along the inlet of the core flowpath and into the high pressure compressor. Baseline conditions were established and parameters were varied to observe accretion characteristics. Data were acquired at altitudes varying from 5 kft to 45 kft, at nominal ice particle Median Volumetric Diameters from 20 µm to 100 µm, and total water contents of 1 g/m3 to 12 g/m3. Metal temperatures were acquired for the inlet guide vane and vane stators 1-2. In-situ measurements of the particle size distribution were acquired upstream and downstream of the engine fan face in order to study particle break-up behavior. Cameras were installed in the engine to capture ice accretions at the leading edge of the fan stator, splitter lip, and inlet guide vane. The goal of this study was to understand the key parameters of accretion, acquire particle break-up data aft of the fan, and generate a unique icing dataset for model development. Significant particle break-up downstream of the fan in the bypass was observed. The metal temperatures on the IGVs and stators show a temperature increase with increasing particle size. Accretion behavior at the fan stator and splitter lip across was very similar. However accretion decreased with increasing particle size at the IGVs.

Error Evaluation in Metal Surface Temperature Prediction Due to Non-Simulation of the Effect of Water Impingement and Runback Water on an Anti-Iced Turboprop Engine Inlet Guide Vane Structure

2008 ◽  
Author(s):  
Javier Castillo ◽  
Gema Ortega
Keyword(s):  
Surface Temperature ◽  
Material Selection ◽  
Guide Vane ◽  
Leading Edge ◽  
Simulation Software ◽  
Inlet Guide Vane ◽  
Development Phase ◽  
Trailing Edge ◽  
Airfoil Surface ◽  
Effect Of Water

During the optimization of the TP400D6 engine (powering the A400M military transport aircraft), the mechanical design of the Front Bearing Structure has proven to be one of the most challenging topics in the engine development programme. One of the leading technical subjects has been the design and optimization of the thermal anti-icing system of the component. When non-specific icing simulation software tools are available, the effect of the water impingement and runback water is difficult to simulate. The objective of this paper is to show one particular aspect learnt during the design and development phase of the project: the evaluation of the error obtained in the calculation of metal temperatures on an antiiced airfoil surface due to the effect of water impingement and runback water. TP400D6 engine front end arrangement consists of a single radial structure after the engine air intake performing both the structural and aerodynamic function, transmitting bearing and high propeller loads and being the compressor IGVs. The anti-icing system employs hot compressor bleed air circulating internally in the component through a series of internal channels and passages and exiting the airfoil through trailing edge holes. Due to airfoil aerodynamic constraints and material selection, it was realised in the earlier stages of the project that it was not possible to heat the whole vane profile up to the trailing edge. In consequence, the effects of the impingement water and runback of non-evaporated water from the intake and the IGV leading edge itself, play a key role on the determination of the airfoil surface temperature and potential ice accretion. Because of not having a specific icing simulation software, water impingement and runback water effects cannot be predicted with sufficient accuracy. During the engine programme’s development phase, a dedicated component antiicing rig test was conducted in order to evaluate and obtain a closer approximation of the real behaviour of the system. The scope of this paper is to go through the details of the aforementioned effect of icing water on airfoil surface temperature, focusing on the discrepancies between predicted temperatures and test rig measured temperatures. Typical thermal modelling is used, which incorporates the best possible understanding of the water particle impingement pattern onto the airfoils and flow lines distribution around the IGV profiles. Results from the rig test have been applied to the traditional thermal model in order to improve the thermal prediction simulation and understanding of the component.

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