scholarly journals Stability assessment of an airflow distorted military engine’s FAN

2017 ◽  
Vol 232 (13) ◽  
pp. 2584-2592 ◽  
Author(s):  
T Triantafyllou ◽  
T Nikolaidis ◽  
M Diakostefanis ◽  
P Pilidis
Keyword(s):  
Total Pressure ◽  
Turbine Engines ◽  
Interface Plane ◽  
Gas Turbine Engines ◽  
Flow Distortion ◽  
Stall Margin ◽  
Surge Margin ◽  
General Dynamics ◽  
Total Temperature ◽  
Model Geometry

Military aircraft are often subjected to severe flight maneuvers with high angles of attack and angles of sideslip. These flight attitudes induce non-uniformity in flow conditions to their gas turbine engines, which may include distortion of inlet total pressure and total temperature at the aerodynamic interface plane. Operation of the downstream engine’s compression system may suffer reduced aerodynamic performance and stall margin, and increased blade stress levels. The present study presents a methodology of evaluating the effect of inlet flow distortion on the engine’s fan stability. The flow distortion examined was induced to the aerodynamic interface plane by means of changing the aircraft’s flight attitude. The study is based on the steady-state flow results from 27 different flight scenarios that have been simulated in computational fluid dynamics. As a baseline model geometry, an airframe inspired by the General Dynamics/LMAERO F-16 aircraft was chosen, which has been exposed to subsonic incoming airflow with varying direction resembling thus different aircraft flight attitudes. The results are focused on the total pressure distribution on the engine’s (aerodynamic interface plane) face and how this is manifested at the operation of the fan. Based on the results, it was concluded that the distorted conditions cause a shift of the surge line on the fan map, with the amount of shift to be directly related to the severity of these distorted conditions. The most severe flight attitude in terms of total pressure distortion, among the tested ones, caused about 7% surge margin depletion comparing to the undistorted value.

Total pressure distortion levels at the aerodynamic interface plane of a military aircraft

2015 ◽  
Vol 119 (1219) ◽  
pp. 1147-1166 ◽  
Author(s):  
T. Triantafyllou ◽  
T. Nikolaidis ◽  
M. Diakostefanis ◽  
P. Pilidis
Keyword(s):  
Total Pressure ◽  
Turbine Engines ◽  
Interface Plane ◽  
Gas Turbine Engines ◽  
Stall Margin ◽  
Average Value ◽  
Circumferential Extent ◽  
Total Temperature ◽  
Computational Fluid Dynamics Cfd ◽  
High Angles Of Attack

AbstractMilitary aircrafts are often subjected to severe flight maneuvers with high Angles-of -Attack (AOA) and Angles of Sideslip (AOSS). These flight attitudes induce non-uniform in flow conditions to their gas turbine engines which may include distortion of inlet total pressure and total temperature at the Aerodynamic Interface Plane (AIP). Operation of the downstream compression system with distorted inflow typically results in reduced aerodynamic performance, reduced stall margin, and increased blade stress levels. In the present study the steady state total pressure distortion induced to the Aerodynamic Interface Plane due to the aircraft’s flight attitude have been estimated in terms of distortion descriptors. The distorted conditions at the interface between the intake and the engine have been predicted by using Computational Fluid Dynamics (CFD), where 33 different aircraft flight attitudes have been tested. Based on the obtained results the effect of Angle-of-Attack (AOA) and Angle of Side Slip (AOSS) on the distortion descriptors have been studied. The results showed that the distortion effect becomes more pronounced whenever this specific airframe configuration is exposed to incoming flow with an AOSS. Among the tested cases, the greatest total pressure defect at the AIP in terms of difference from the average value and of circumferential extent was calculated for the flight attitudes of 0·35M flight with 0° AOA and 8° AOSS and 0·35M fight with 16° AOA and 16° AOSS.

scholarly journals JET ENGINE INLET DISTORTION SCREEN AND DESCRIPTOR EVALUATION

2017 ◽  
Vol 57 (1) ◽  
pp. 22-31 ◽  
Author(s):  
Jiří Pečinka ◽  
Gabriel Thomas Bugajski ◽  
Petr Kmoch ◽  
Adolf Jílek
Keyword(s):  
Total Pressure ◽  
Basic Flow ◽  
Interface Plane ◽  
Inlet Channel ◽  
Flow Distortion ◽  
Turbine Engine ◽  
Inlet Distortion ◽  
Mass Flow Rates ◽  
Total Temperature ◽  
The University

Total pressure distortion is one of the three basic flow distortions (total pressure, total temperature and swirl distortion) that might appear at the inlet of a gas turbine engine (GTE) during operation. Different numerical parameters are used for assessing the total pressure distortion intensity and extent. These summary descriptors are based on the distribution of total pressure in the aerodynamic interface plane. There are two descriptors largely spread around the world, however, three or four others are still in use and can be found in current references. The staff at the University of Defence decided to compare the most common descriptors using basic flow distortion patterns in order to select the most appropriate descriptor for future department research. The most common descriptors were identified based on their prevalence in widely accessible publications. The construction and use of these descriptors are reviewed in the paper. Subsequently, they are applied to radial, angular, and combined distortion patterns of different intensities and with varied mass flow rates. The tests were performed on a specially designed test bench using an electrically driven standalone industrial centrifugal compressor, sucking air through the inlet of a TJ100 small turbojet engine. Distortion screens were placed into the inlet channel to create the desired total pressure distortions. Of the three basic distortions, only the total pressure distortion descriptors were evaluated. However, both total and static pressures were collected using a multi probe rotational measurement system.

Component Integration and Loss Sources in 3 – 5 kW Gas Turbines

2005 ◽  
Author(s):  
M. A. Monroe ◽  
A. H. Epstein ◽  
H. Kumakura ◽  
K. Isomura
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Large Scale ◽  
Engine Performance ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Flow Distortion ◽  
Thermal Isolation ◽  
Component Integration ◽  
Measured Efficiency

The performance of a regenerated gas turbine generator in the 3–5 kW power range has been analyzed to understand why its measured efficiency was on the order of 6% rather than the 20% suggested by consideration of its components’ efficiencies as measured on rigs. This research suggests that this discrepancy can be primarily attributed to heat and fluid leaks not normally considered in the analysis of large gas turbine engines because they are not as important at large scale. In particular, fluid leaks among the components and heat leakage from the hot section into the compressor flow path contributed the largest debits to the engine performance. Such factors can become more important as the engine size is reduced. Other non-ideal effects reducing engine performance include temperature flow distortion at the entrance to both the compressor and turbine. A cycle calculation including all of the above effects matched measured engine data. It suggests that relatively simple changes such as thermal isolation and leak sealing can increase both power output and efficiency of this engine, over 225% in the latter case. The validity of this analysis was demonstrated on an engine in which partial thermal isolation and improved sealing resulted in a more than 40% increase in engine output power.

scholarly journals Flight Research Using F100 Engine P680063 in the NASA F-15 Airplane

1995 ◽  
Author(s):  
Frank W. Burcham ◽  
Timothy R. Conners ◽  
Michael D. Maxwell
Keyword(s):  
Control System ◽  
Flight Control ◽  
Flight Test ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Stall Margin ◽  
Engine Control System ◽  
Engine Thrust ◽  
Research Activities ◽  
Margin Control

The value of flight research in developing and evaluating gas turbine engines is high. NASA Dryden Flight Research Center has been conducting flight research on propulsion systems for many years. The F100 engine has been tested in the NASA F-15 research airplane in the last three decades. One engine in particular, S/N P680063, has been used for the entire program and has been flown in many pioneering propulsion flight research activities. Included are detailed flight-to-ground facility tests; tests of the first production digital engine control system, the first active stall margin control system, the first performance-seeking control system; and the first use of computer-controlled engine thrust for emergency flight control. The flight research has been supplemented with altitude facility tests at key times. This paper presents a review of the tests of engine P680063, the F-15 airplanes in which it flew, and the role of the flight test in maturing propulsion technology.

High-temperature high-frequency turbine exit flow field measurements in a military engine with a cooled unsteady total pressure probe

2011 ◽  
Vol 225 (7) ◽  
pp. 954-963 ◽  
Author(s):  
M Mersinligil ◽  
J Desset ◽  
J F Brouckaert
Keyword(s):  
High Temperature ◽  
Total Pressure ◽  
Pressure Sensors ◽  
Field Measurements ◽  
Fast Response ◽  
Turbine Engines ◽  
Pressure Probe ◽  
Gas Turbine Engines ◽  
Time Resolved ◽  
High Bandwidth

The measurement of unsteady pressures within the hot components of gas turbine engines still remains a true challenge for test engineers. Several high-temperature pressure sensors have been developed, but so far, their applications are restricted to unsteady wall static pressure measurements. Because of the severe flow conditions such as turbine inlet temperatures of 1700 °C and pressures of 50 bar or more in the most advanced aero-engine designs, few (if any) experimental techniques exist to measure the time-resolved flow total pressure inside the gas path. This article describes the measurements performed at the turbine exit of a military engine with a cooled fast response total pressure probe. The probe concept is based on the use of a conventional miniature piezo-resistive pressure sensor, located in the probe tip to achieve a bandwidth of at least 40 kHz. Due to the extremely harsh conditions, the probe and sensor are heavily water cooled. The probe was designed to be continuously immersed into the hot gas stream to obtain time series of pressure with a high bandwidth and therefore statistically representative average fluctuations at the blade passing frequency (BPV). The experimental results obtained with a second-generation prototype are presented. The probe was immersed into the engine through the bypass duct between turbine exit and flame-holders of the afterburner of a Volvo RM12 engine, at exhaust temperatures above 900 °C. The probe was able to resolve the BPV (∼17 kHz) and several harmonics up to 100 kHz.

Wind-tunnel tests of the ERICA tiltrotor optimised air-intake

2018 ◽  
Vol 122 (1251) ◽  
pp. 821-837
Author(s):  
G. Gibertini ◽  
A. Zanotti ◽  
G. Campanardi ◽  
F. Auteri ◽  
D. Zagaglia ◽  
...  
Keyword(s):  
Wind Tunnel ◽  
Total Pressure ◽  
Internal Surface ◽  
Scale Model ◽  
Interface Plane ◽  
Shape Optimisation ◽  
Duct Flow ◽  
Flow Distortion ◽  
Wind Tunnel Tests ◽  
Air Intake

ABSTRACTWind-tunnel tests were carried out to evaluate the performance of the Computational Fluid Dynamics (CFD)-based air-intake duct shape optimisation of the European platform tiltrotor ERICA. A 1/2.5 scale model including the nacelle, the external portion of the wing and two interchangeable internal ducts reproducing the baseline and optimised shape were manufactered to be tested in the large wind tunnel of Politecnico di Milano. Moreover, tests were carried out with the model equipped with rotating blade stubs. The comprehensive experimental campaign included tests reproducing different forward flight conditions of the aircraft including cruise and conversion phases. The evaluation of the internal duct performance was carried out by measuring total pressure losses and flow distortion by directional probes at the Aerodynamic Interface Plane (AIP). Additional pressure measurements were carried out on the internal surface of the duct to compare the pressure distributions along the air-intake. The experimental results confirmed that the optimised duct offers significantly improved performance with respect to the baseline configuration not only in cruise, representing the flight condition considered for the CFD optimisation, but also for the conversion condition. In particular, a remarkable reduction of the total pressure drop at the AIP was found with the optimised duct with the only exception for the stubs-on configuration in cruise. Indeed, the present investigation highlighted that the design of the blade stubs, particularly their length, represents a very critical aspect for air-intake performance tests due to significant disturbances that could be induced by the stubs’ wake on the internal duct flow.

scholarly journals Design of an Economical Counter Rotating Fan: Comparison of the Calculated and Measured Steady and Unsteady Results

2012 ◽  
Author(s):  
Timea Lengyel-Kampmann ◽  
Andreas Bischoff ◽  
Robert Meyer ◽  
Eberhard Nicke
Keyword(s):  
High Resolution ◽  
Total Pressure ◽  
High Efficiency ◽  
Optimization Method ◽  
Test Facility ◽  
Hot Wire ◽  
Stall Margin ◽  
Low Speed ◽  
Total Temperature ◽  
German Aerospace

Within the framework of the EU funded Project VITAL, SNECMA (Group Safran), as the work package leader, developed a counter rotating low-speed fan-concept for a high bypass ratio engine. The detailed aerodynamic and mechanical optimization of one blading version (CRTF2.b) was carried out at the German Aerospace Center (DLR), by applying one of the newest design methods featuring a multi-objective automatic optimization method based on an Evolutionary Algorithm [1]. The final design goals were high efficiency, a sufficient stall margin and adequate acoustic performances for the given cycle parameters. The fan stage developed was tested in an anechoic test facility at CIAM in Moscow. The test routine included the measurement of the performance map based on total pressure and total temperature measurements at the inlet and the outlet of the test rig and acoustic measurement as well. The unsteady flow field of the low speed Contra-Rotating Turbo Fan has been measured with four hot-wire probes at different axial positions. In the evaluation the measured data are compared with high resolution CFD results. Special emphasis was given to the comparison of the radial distribution of total pressure and total temperature in the bypass channel, the comparison of the measured and the calculated fan maps and to the comparison of the hot-wire measurements with high resolution, unsteady CFD results. The tests and the URANS-results confirmed the design goals.

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