Engine Ground Start Characteristics of Fighter Aircraft at High Altitude Airbase

2019 ◽  
Author(s):  
Santhosh Kasram ◽  
Sajath Kumar Manoharan ◽  
Mahesh P. Padwale ◽  
G. P. Ravishankar
Keyword(s):  
High Altitude ◽  
Gas Turbine ◽  
Weather Conditions ◽  
Gas Turbine Engine ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Alternative Methods ◽  
Cold Weather ◽  
Oil Viscosity ◽  
Turbine Engine

Abstract The challenges faced during starting of an aircraft gas turbine engine using a Jet Fuel Starter (JFS) at high altitude airbase are discussed in this paper. Autonomous ground starts at high altitude airbase in soaked sub-zero temperature condition without any external ground support assistance is a challenge. Generally, the start cycle (sub-idle speed) at sub-zero temperatures of a gas turbine engine at high altitudes is influenced by several factors. Drag loads are estimated due to change in lube oil viscosity of engine gearbox and accessory gear box that affects available torque margin of a starter. These estimated loads are superimposed on starter characteristics to identify the available margins for successful starts. The cold start is particularly severe, since it increases the tip clearance between rotor and casing of the engine due to difference in its thermal growth. Higher tip clearances significantly degrade compressor surge margin and results in rotating stall. Inconsistent engine starts were resolved by adopting alternative methods without any change in hardware. This paper presents set of methods used to overcome inconsistent engine starts at high altitude cold weather conditions.

The Effect of Heat Transfer Coefficient Increase on Tip Clearance Control in H.P. Compressors in Gas Turbine Engine

2013 ◽  
Author(s):  
Godwin Ita Ekong ◽  
Christopher A. Long ◽  
Peter R. N. Childs
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Time Constant ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Test Facility ◽  
Turbine Engine

Compressor tip clearance for a gas turbine engine application is the radial gap between the stationary compressor casing and the rotating blades. The gap varies significantly during different operating conditions of the engine due to centrifugal forces on the rotor and differential thermal expansions in the discs and casing. The tip clearance in the axial flow compressor of modern commercial civil aero-engines is of significance in terms of both mechanical integrity and performance. In general, the clearance is of critical importance to civil airline operators and their customers alike because as the clearance between the compressor blade tips and the casing increases, the aerodynamic efficiency will decrease and therefore the specific fuel consumption and operating costs will increase. This paper reports on the development of a range of concepts and their evaluation for the reduction and control of tip clearance in H.P. compressors using an enhanced heat transfer coefficient approach. This would lead to improvement in cruise tip clearances. A test facility has been developed for the study at the University of Sussex, incorporating a rotor and an inner shaft scaled down from a Rolls-Royce Trent aero-engine to a ratio of 0.7:1 with a rotational speed of up to 10000 rpm. The idle and maximum take-off conditions in the square cycle correspond to in-cavity rotational Reynolds numbers of 3.1×106 ≤ Reφ ≤ 1.0×107. The project involved modelling of the experimental facilities, to demonstrate proof of concept. The analysis shows that increasing the thermal response of the high pressure compressor (HPC) drum of a gas turbine engine assembly will reduce the drum time constant, thereby reducing the re-slam characteristics of the drum causing a reduction in the cold build clearance (CBC), and hence the reduction in cruise clearance. A further reduction can be achieved by introducing radial inflow into the drum cavity to further increase the disc heat transfer coefficient in the cavity; hence a further reduction in disc drum time constant.

scholarly journals Gemini T-20G-8 XM1 Tank APU

1981 ◽  
Author(s):  
C. Rodgers ◽  
J. Zeno ◽  
E. A. Drury ◽  
A. Karchon
Keyword(s):  
Gas Turbine ◽  
Power Unit ◽  
Weather Conditions ◽  
Cold Weather ◽  
Turbine Engine ◽  
Auxiliary Power Unit ◽  
Auxiliary Power ◽  
Generator Sets ◽  
Main Engine ◽  
The U.S

Auxiliary power is often provided on combat vehicles in the U.S. Army for battery charging, operation of auxiliary vehicle equipment when the main engine is not running, or to provide assistance in starting the main engine in extreme cold weather conditions. The use of a gas turbine for these applications is particularly attractive, due to its small size and lightweight. In November 1978, the U.S. Army Tank-Automotive Research and Development Command, Warren, MI awarded a contract to the Turbomach Division of Solar Turbines International, San Diego, CA, for the development of a 10 kW 28 vdc gas turbine powered auxiliary power unit (APU) for installation in the XM1 main battle tank. This paper describes the general features of the Solar Turbomach T-20G-8 Auxiliary Power Unit, a single-shaft gas turbine driven generator set which has been developed under this contract. This APU is one of the family of Gemini powered APUs and is a derivative of the U.S. Army 10 kW gas turbine engine-driven, 60 and 400 Hz generator sets developed by Solar. The electrical components were newly developed for this particular application. Currently, the APU is in qualification testing both in the laboratory and in the XM1 main battle tank.

Diagnosing Propulsion Systems of a Vessels in Varying Marine Conditions

2013 ◽  
Vol 199 ◽  
pp. 9-14
Author(s):  
Adam Charchalis
Keyword(s):  
Gas Turbine ◽  
Weather Conditions ◽  
Gas Turbine Engine ◽  
Propulsion Systems ◽  
Turbine Engine ◽  
External Conditions ◽  
The Way

The paper presents some problems of carrying out measurements of energetic characteristics and vessels performance in the conditions of sea examinations. We present the influence of external conditions in the change of vessels hull resistance and propeller characteristics as well as the influence of weather conditions in the results of examinations and characteristics of gas turbine engine. We also discuss the manner of reducing the results of measurements to the standard conditions. We present the way of preparing propulsion characteristics and the analysis of examination uncertainty for the measurement of torque.

Qualification of a Small Gas Turbine Engine as a Starter Unit

2017 ◽  
Vol 0 (0) ◽  
Author(s):  
R. K. Mishra ◽  
Prashant Kumar ◽  
K. S. Jayasihma ◽  
S. N. Mistry
Keyword(s):  
High Altitude ◽  
Gas Turbine ◽  
Sea Level ◽  
Rotational Speed ◽  
Gas Turbine Engine ◽  
Point Of View ◽  
High Rotational Speed ◽  
Turbine Engine ◽  
Starter Unit

AbstractThe certification philosophy plays an important role in ensuring the airworthiness qualification of a small gas turbine engine designed as a starter unit. The small and compactness of the engine, high rotational speed of the rotors, requirements of torque and starting capability at sea level to high altitude airfields and consecutive starts within stipulated time impose a great challenge from airworthiness point of view. This paper presents the methodology adopted and various stages of qualification, standards followed and results based on which the starter unit has been qualified for fitment on the designated aircraft.

Implanted Component Faults and Their Effects on Gas Turbine Engine Performance

1992 ◽  
Vol 114 (2) ◽  
pp. 174-179 ◽  
Author(s):  
J. D. MacLeod ◽  
V. Taylor ◽  
J. C. G. Laflamme
Keyword(s):  
Gas Turbine ◽  
National Research Council ◽  
Engine Performance ◽  
Turbine Rotor ◽  
Gas Turbine Engine ◽  
Tip Clearance ◽  
Turbine Engine ◽  
Turbine Rotor Blade ◽  
Department Of National Defence ◽  
Project Objectives

Under the sponsorship of the Canadian Department of National Defence, the Engine Laboratory of the National Research Council of Canada (NRCC) has established a program for the evaluation of component deterioration on gas turbine engine performance. The effect is aimed at investigating the effects of typical in-service faults on the performance characteristics of each individual engine component. The objective of the program is the development of a generalized fault library, which will be used with fault identification techniques in the field, to reduce unscheduled maintenance. To evaluate the effects of implanted faults on the performance of a single spool engine, such as an Allison T56 turboprop engine, a series of faulted parts were installed. For this paper the following faults were analyzed: (a) first-stage turbine nozzle erosion damage; (b) first-stage turbine rotor blade untwist; (c) compressor seal wear; (d) first and second-stage compressor blade tip clearance increase. This paper describes the project objectives, the experimental installation, and the results of the fault implantation on engine performance. Discussed are performance variations on both engine and component characteristics. As the performance changes were significant, a rigorous measurement uncertainty analysis is included.

Probabilistic High-Cycle Fretting Fatigue Assessment of Gas Turbine Engine Components

2011 ◽  
Author(s):  
Kwai S. Chan ◽  
Michael P. Enright ◽  
Patrick J. Golden ◽  
Samir Naboulsi ◽  
Ramesh Chandra ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Turbine Blades ◽  
Real Life ◽  
Fretting Fatigue ◽  
Gas Turbine Engine ◽  
Rotating Stall ◽  
Worst Case ◽  
Design Assessment ◽  
Turbine Engine ◽  
Engine Components

High-cycle fatigue (HCF) is arguably one of the costliest sources of in-service damage in military aircraft engines. HCF of turbine blades and disks can pose a significant engine risk because fatigue failure can result from resonant vibratory stresses sustained over a relatively short time. A common approach to mitigate HCF risk is to avoid dangerous resonant vibration modes (first bending and torsion modes, etc.) and instabilities (flutter and rotating stall) in the operating range. However, it might be impossible to avoid all the resonance for all flight conditions. In this paper, a methodology is presented to assess the influences of HCF loading on the fracture risk of gas turbine engine components subjected to fretting fatigue. The methodology is based on an integration of a global finite element analysis of the disk-blade assembly, numerical solution of the singular integral equations using the CAPRI (Contact Analysis for Profiles of Random Indenters) and Worst Case Fret methods, and risk assessment using the DARWIN (Design Assessment of Reliability with Inspection) probabilistic fracture mechanics code. The methodology is illustrated for an actual military engine disk under real life loading conditions.

scholarly journals A Simplified Gas Turbine Engine Model With Heat Storage/Tip Clearance Effects

1993 ◽  
Author(s):  
T. H. Wong
Keyword(s):  
Gas Turbine ◽  
Heat Storage ◽  
Gas Turbine Engine ◽  
Tip Clearance ◽  
Metal Temperature ◽  
Transient Method ◽  
Transient Responses ◽  
Turbine Engine ◽  
Engine Simulation ◽  
Engine Model

A simplified turboshaft gas turbine engine model called the direct transient method (DTM) model has been developed. The DTM model consists of table look-up data generated from the actual engine data or the transient engine simulation. The DTM model accounts for heat storage, tip clearance and volume dynamics effects. It can, therefore, better predict engine transient responses and turbine metal temperature than the traditional engine horsepower extraction (HPX) model. This paper presents in detail the DTM methodology for generating accurate simplified engine models of transient performance. Comparisons of engine transient responses between the DTM and HPX models are provided.

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