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 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.

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.

Analysis of Turbofan Performance Under Total Pressure Distortion at Various Operating Points

2015 ◽  
Author(s):  
David B. Weston ◽  
Steven E. Gorrell ◽  
Matthew L. Marshall ◽  
Carol V. Wallis
Keyword(s):  
Total Pressure ◽  
Operating Conditions ◽  
Rotor Blades ◽  
Cfd Simulations ◽  
Blade Loading ◽  
Distortion Analysis ◽  
Work Input ◽  
Total Temperature ◽  
Computational Fluid Dynamics Cfd ◽  
Operating Points

Inlet distortion is an important consideration in fan performance. The focus of this paper is a series of high-fidelity time accurate Computational Fluid Dynamics (CFD) simulations of a multistage fan at choke, design, and near stall operating conditions. These investigate distortion transfer and generation as well as the underlying flow physics of these phenomena under different operating conditions. The simulations are performed on the full annulus of a 3 stage fan and are analyzed. The code used to carry out these simulations is a modified version of OVERFLOW 2.2. The inlet is specified as a 1/rev total pressure distortion. Analysis includes the phase and amplitude of total temperature and pressure distortion through each stage of the fan and blade loading. The total pressure distortion does not change in severity through the fan, but the peak pressure distortion rotates by as much as 45° at the near stall point. This is due to a variation in the work input around the blades of the rotor. This variation is also responsible for the generation of total temperature distortion in the fan. The rotation of the total temperature distortion becomes more pronounced as the fan approaches stall, and the total temperature distortion levels increase. The amount of work performed by a single blade can vary by as much as 25% in the first stage at near stall. The variation in work becomes more pronounced as the fan approaches stall. The passage shock in the rotor blades moves nearly 20% of the blade chord in both the peak efficiency and near stall cases.

An Experimental Investigation Into the Impacts of Varying the Circumferential Extent of Tip-Low Total Pressure Distortion on Fan Stability

2021 ◽  
Author(s):  
Oliver Allen ◽  
Alejandro Castillo Pardo ◽  
Cesare A. Hall
Keyword(s):  
Total Pressure ◽  
Stability Margin ◽  
Jet Engines ◽  
Stall Inception ◽  
Stall Margin ◽  
High Incidence ◽  
Circumferential Extent ◽  
Low Distortion ◽  
The Stability ◽  
The Impact

Abstract Future jet engines with shorter and thinner intakes present a greater risk of intake separation. This leads to a complex tip-low total pressure distortion pattern of varying circumferential extent. In this paper, an experimental study has been completed to determine the impact of such distortion patterns on the operating range and stalling behaviour of a low-speed fan rig. Unsteady casing static pressure measurements have been made during stall events in 11 circumferential extents of tip-low distortion. The performance has been measured and detailed area traverses have been performed at rotor inlet and outlet in 3 of these cases — clean, axisymmetric tip-low and half-annulus tip-low distortion. Axisymmetric tip-low distortion is found to reduce stall margin by 8%. It does not change the stalling mechanism compared to clean inflow. In both cases, high incidence at the tip combined with growth of the casing boundary layer drive instability. In contrast, half-annulus tip-low distortion is found to reduce stall margin by only 4% through a different mechanism. The distortion causes disturbances in the measured casing pressure signals to grow circumferentially in regions of high incidence. Stall occurs when these disturbances do not decay fully in the undistorted region. As the extent of the distorted sector is increased, the stability margin is found to reduce continuously. However, the maximum disturbance size before stall inception is found to occur at intermediate values of distorted sector extent. This corresponds to distortion patterns that provide sufficient circumferential length of undistorted region for disturbances to decay fully before they return to the distorted sector. It is found that as the extent of the tip-low distortion sector is increased, the circumferential size of the stall cell that develops is reduced. However, its speed is found to remain approximately constant at 50% of the rotor blade speed.

Evaluation of a 3D Viscous Code for Turbomachinery Flows

1997 ◽  
Author(s):  
W. John Calvert ◽  
Andrew W. Stapleton ◽  
Paul R. Emmerson ◽  
Cecil R. Buchanan ◽  
Christopher M. Nott
Keyword(s):  
Boundary Layer ◽  
Turbulent Flows ◽  
Strong Shock ◽  
Turbulence Models ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Complex Flows ◽  
Analysis And Design ◽  
Computational Fluid Dynamics Cfd ◽  
Wave Boundary

Computational Fluid Dynamics (CFD) codes play an increasingly important role in the design and development of turbomachinery for modern gas turbine engines. As a result additional emphasis is being placed on the evaluation of the codes to ensure that they are working correctly and to indicate the accuracy which is likely to be achieved in practice. At DERA a programme of work has been carried out to evaluate the TRANSCode 3D viscous flow code, which was developed from the BTOB3D solver written by Dawes in 1986. A three part strategy for the validation and calibration of the code was adopted, covering comparisons with boundary layer test cases, Q3D compressor cascades and full 3D cases. The results indicated that the grids currently employed for turbomachinery flows limit the accuracy achieved for cases where there is significant laminar flow. For turbulent flows the Baldwin-Lomax turbulence model gives reasonably accurate results for 2D flows in near equilibrium, but it is less satisfactory for more complex flows, when the concept of a simple 2D boundary layer does not apply, and for strong shock wave/boundary layer interactions. Overall it is considered that the code is a valuable tool for turbomachinery analysis and design, but solutions must be assessed with care. Alternative turbulence models and other developments are being pursued for future versions of the code.

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.

High-Fidelity Numerical Analysis of Per-Rev-Type Inlet Distortion Transfer in Multistage Fans—Part II: Entire Component Simulation and Investigation

2010 ◽  
Vol 132 (4) ◽  
Author(s):  
Jixian Yao ◽  
Steven E. Gorrell ◽  
Aspi R. Wadia
Keyword(s):  
Total Pressure ◽  
Navier Stokes ◽  
High Fidelity ◽  
Inlet Distortion ◽  
Unsteady Rans ◽  
Total Temperature ◽  
Computational Fluid Dynamics Cfd ◽  
Last Stage ◽  
Remarkable Agreement ◽  
Engine Development

Part I of this paper validated the ability of the unsteady Reynolds-Averaged Navier-Stokes (RANS) solver PTURBO to accurately simulate distortion transfer and generation through selected blade rows of two multistage fans. In this part, unsteady RANS calculations were successfully applied to predict the 1/rev inlet total pressure distortion transfer in the entirety of two differently designed multistage fans. This paper demonstrates that high-fidelity computational fluid dynamics (CFD) can be used early in the design process for verification purposes before hardware is built and can be used to reduce the number of distortion tests, hence reducing engine development cost. The unsteady RANS code PTURBO demonstrated remarkable agreement with the data, accurately capturing both the magnitude and the profile of total pressure and total temperature measurements. Detailed analysis of the flow physics identified from the CFD results has led to a thorough understanding of the total temperature distortion generation and transfer mechanism, especially for the spatial phase difference of total pressure and total temperature profiles. The analysis illustrates that the static parameters are more revealing than their stagnation counterpart and that pressure and temperature rise are more revealing while the pressure and temperature ratio could be misleading. The last stage is effectively throttled by the inlet distortion even though the overall engine throttle remains unchanged. The total temperature distortion generally grows as flow passes through the fan stages.

Three-dimensional unsteady investigation of HP turbine stages

2006 ◽  
Vol 220 (2) ◽  
pp. 155-167 ◽  
Author(s):  
P Adami ◽  
F Martelli
Keyword(s):  
Pressure Fluctuations ◽  
Three Dimensional ◽  
Numerical Approach ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Turbulent Closure ◽  
Rotor Stator Interaction ◽  
Computational Fluid Dynamics Cfd ◽  
Finite Volume Approach ◽  
Unsteady Numerical Simulation

This article deals with a three-dimensional unsteady numerical simulation of the unsteady rotor—stator interaction in a HP turbine stage. The numerical approach consists of a computational fluid dynamics (CFD) parallel code, based on an upwind total variation diminishing finite volume approach. The computation has been carried out using a sliding plane approach with hybrid unstructured meshes and a two-equation turbulent closure. The turbine rig under investigation is representative of the first stage of aeronautic gas turbine engines. A brief description of the cascade, the experimental setup, and the measuring technique is provided. Time accurate CFD computations of pressure fluctuations and Nusselt number are discussed against the experimental data.

A Simple Physics-Based Model for Particle Rebound and Deposition in Turbomachinery

2017 ◽  
Vol 139 (8) ◽  
Author(s):  
J. P. Bons ◽  
R. Prenter ◽  
S. Whitaker
Keyword(s):  
Gas Turbine ◽  
Particle Deposition ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Cfd Simulations ◽  
Model Predictions ◽  
Computational Fluid Dynamics Cfd ◽  
Parametric Dependencies ◽  
Key Aspects ◽  
Physical Phenomena

A new model is proposed for predicting particle rebound and deposition in environments relevant to gas turbine engines. The model includes the following physical phenomena: elastic deformation, plastic deformation, adhesion, and shear removal. It also incorporates material property sensitivity to temperature and tangential-normal rebound velocity cross-dependencies observed in experiments. The model is well-suited for incorporation in computational fluid dynamics (CFD) simulations of complex gas turbine flows due to its algebraic (explicit) formulation. Model predictions are compared to coefficient of restitution data available in the open literature as well as deposition results from two different high-temperature turbine deposition facilities. While the model comparisons with experiments are in many cases promising, several key aspects of particle deposition remain elusive. The simple phenomenological nature of the model allows for parametric dependencies to be evaluated in a straightforward manner. It is hoped that this feature of the model will aid in identifying and resolving the remaining stubborn holdouts that prevent a universal model for particle deposition.

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