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.

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 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%.

Mechanical Design and Testing of a 2.5 MW sCO2 Compressor Loop

2021 ◽  
Author(s):  
Stefan D. Cich ◽  
J. Jeffrey Moore ◽  
Chris Kulhanek ◽  
Meera Day Towler ◽  
Jason Mortzheim
Keyword(s):  
High Efficiency ◽  
Mechanical Design ◽  
Guide Vane ◽  
Operating Conditions ◽  
Inlet Guide Vane ◽  
Inlet Temperature ◽  
Design Changes ◽  
Wide Range ◽  
Inlet Conditions ◽  
Design And Testing

Abstract An enabling technology for a successful deployment of the sCO2 close-loop recompression Brayton cycle is the development of a compressor that can maintain high efficiency for a wide range of inlet conditions due to large variation in properties of CO2 operating near its dome. One solution is to develop an internal actuated variable Inlet Guide Vane (IGV) system that can maintain high efficiency in the main and re-compressor with varying inlet temperature. A compressor for this system has recently been manufactured and tested at various operating conditions to determine its compression efficiency. This compressor was developed with funding from the US DOE Apollo program and industry partners. This paper will focus on the design and testing of the main compressor operating near the CO2 dome. It will look at design challenges that went into some of the decisions for rotor and case construction and how that can affect the mechanical and aerodynamic performance of the compressor. This paper will also go into results from testing at the various operating conditions and how the change in density of CO2 affected rotordynamics and overall performance of the machine. Results will be compared to expected performance and how design changes were implanted to properly counter challenges during testing.

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.

Aerothermal Performance of Shielded Vane Design

2017 ◽  
Vol 139 (11) ◽  
Author(s):  
Ioanna Aslanidou ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Total Pressure ◽  
High Speed ◽  
Guide Vane ◽  
Leading Edge ◽  
Experimental Facility ◽  
Numerical Studies ◽  
Nozzle Guide Vane ◽  
Inlet Conditions ◽  
Nozzle Guide

This paper presents an experimental investigation of the concept of using the combustor transition duct wall to shield the nozzle guide vane leading edge. The new vane is tested in a high-speed experimental facility, demonstrating the improved aerodynamic and thermal performance of the shielded vane. The new design is shown to have a lower average total pressure loss than the original vane, and the heat transfer on the vane surface is overall reduced. The peak heat transfer on the vane leading edge–endwall junction is moved further upstream, to a region that can be effectively cooled as shown in previously published numerical studies. Experimental results under engine-representative inlet conditions showed that the better performance of the shielded vane is maintained under a variety of inlet conditions.

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.

Online Performance Monitoring of Industrial Compressors Using Meanline Modelling

2014 ◽  
Author(s):  
Matteo Cicciotti ◽  
Dionysios P. Xenos ◽  
Ala E. F. Bouaswaig ◽  
Nina F. Thornhill ◽  
Ricardo F. Martinez-Botas
Keyword(s):  
Performance Monitoring ◽  
Guide Vane ◽  
Pressure Ratio ◽  
Inlet Guide Vane ◽  
Mechanical Degradation ◽  
Adaptive Parameters ◽  
Online Performance Monitoring ◽  
Constrained Minimization Problem ◽  
Inlet Conditions ◽  
Process Measurements

This paper proposes a framework for detecting mechanical degradation online and assessing its effect on the performance of industrial compressors. It consists of a model of the machine in undegraded condition and of a degradation adaptive model. The proposed methodology for online degradation detection differentiates itself from those found in the literature as the undegraded model is not linearized and ambient/inlet conditions are explicitly taken into account. The degradation is modelled through adaptive parameters which are estimated and updated online through the solution of a constrained minimization problem within a moving window. It uses available process measurements of flow, pressures, temperatures and composition. The update of the parameters guarantees the model accuracy and it permits the estimation of the effects of mechanical degradation away from the compressor running line. The performance monitoring framework has been successfully applied on an industrial air centrifugal compressor. It was found that after 3250 hours of operation from the previous maintenance the efficiency and the pressure ratio had dropped approximately 5.5% and 2.5% of their respective undegraded values. Furthermore, it was found that the performance deviations from the baseline depend from the position of the operative point in the performance map. In fact, the pressure ratio drop was lower (2%) and efficiency drop was higher (6%) for lower inlet guide vanes opening whereas pressure ratio drop was higher (3%) and efficiency drop was lower (1.6%) for higher inlet guide vane opening.

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.

Leading edge film cooling enhancement for an inlet guide vane with fan-shaped holes

2010 ◽  
Vol 19 (6) ◽  
pp. 514-518 ◽  
Author(s):  
Jian-Jun Liu ◽  
Bai-Tao An ◽  
Jie Liu ◽  
W. Zhan
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Inlet Guide Vane ◽  
Cooling Enhancement

Experimental and Numerical Researches in a Four-Stage Low Speed Research Compressor Facility

2020 ◽  
Author(s):  
Mingmin Zhu ◽  
Xiaoqing Qiang ◽  
Zhenzhou Ju ◽  
Yuchen Ma ◽  
Jinfang Teng
Keyword(s):  
Guide Vane ◽  
Axial Compressor ◽  
Internal Flow ◽  
Flow Fields ◽  
Inlet Guide Vane ◽  
External Diameter ◽  
Low Speed ◽  
Single Passage ◽  
Overall Performance ◽  
Inlet Conditions

Abstract The flow fields in rear stages of multi-stage axial compressor is difficult to measure in detail owing to the limited height and space. Thus, low speed research compressor (LSRC) facilities which are modelled from rear stages have been widely used to explore the internal flow fields and improve compressor design. A newly-designed vertical LSRC facility is established and put into used in Shanghai Jiao Tong University. The construction and design features of this LSRC facility are introduced in this paper. A cantilevered stage has been tested in this test rig. Compressor performance, inter-stage parameters distributions and contours are measured at design point and near stall point. Steady single passage simulations for four-stage compressor are carried out to validate numerical methods and further interpret the internal flow fields in test stages. This vertical LSRC facility consists of inlet guide vane (IGV) and four repeated stages with an external diameter of 1.5 meter and a rotating speed of 900 RPM. The third stage is the mainly tested one, while the first and second stages provide the inlet conditions and the fourth stage provides the outlet conditions. Complete measuring methods and systems are established for this newly-built LRSC facility. The measurements of overall performance and inter-stage flow fields are carried out for test stage with cantilevered stator rows. The simulation for four-stage compressor are also performed for cantilevered configuration. The results of steady single-passage simulation have a similar trend with experimental ones, in terms of overall performance and parameters distributions.

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