Improving the Measurement Uncertainty of Altitude Test Facility for Gas Turbine Engines

2002 ◽  
Vol 26 (11) ◽  
pp. 1496-1502 ◽  
Keyword(s):  
Measurement Uncertainty ◽  
Gas Turbine ◽  
Test Facility ◽  
Turbine Engines ◽  
Gas Turbine Engines

Evaluation of Heat-Resistant Alloys for Marine Gas Turbine Applications

1967 ◽  
Vol 89 (1) ◽  
pp. 23-27 ◽  
Author(s):  
L. J. Fiedler ◽  
R. M. N. Pelloux
Keyword(s):  
Gas Turbine ◽  
Protective Coatings ◽  
Test Facility ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
X Ray Diffraction ◽  
Combustion Gas ◽  
Turbine Engine ◽  
Gas Atmosphere ◽  
Combustion Residue

Materials for the turbine and combustor sections of gas turbine engines were evaluated for their resistance to sulfidation corrosion. The basic evaluation was conducted in a test facility by exposing the materials to a combustion gas atmosphere which simulates conditions of gas composition, corrosive combustion residue, gas velocity, and temperature that are encountered while operating a gas turbine engine in a marine environment. The influence of alloy composition, protective coatings, salt ingestion rates, and fuel sulfur content is discussed in relation to the degree of sulfidation corrosion. The mechanism of sulfidation corrosion attack, as determined by electron microprobe analyses and X-ray diffraction studies of corroded materials, is also discussed.

Advanced Multi-Cup Fuel Injector Technology for Environmentally Responsible Aviation Gas Turbine Engines

2015 ◽  
Author(s):  
Erlendur Steinthorsson ◽  
Adel Mansour ◽  
Brian Hollon ◽  
Michael Teter ◽  
Clarence Chang
Keyword(s):  
Gas Turbine ◽  
Direct Injection ◽  
Pressure Ratio ◽  
Test Facility ◽  
Turbine Engines ◽  
Practical Implementation ◽  
Gas Turbine Engines ◽  
Fuel Injector ◽  
Ambient Conditions ◽  
Environmentally Responsible

Participating in NASA’s Environmentally Responsible Aviation (ERA) Project, Parker Hannifin built and tested multipoint Lean Direct Injection (LDI) fuel injectors designed for NASA’s N+2 55:1 Overall Pressure-Ratio (OPR) gas turbine engine cycles. The injectors are based on Parker’s earlier three-zone injector (3ZI) which was conceived to enable practical implementation of multipoint LDI schemes in conventional aviation gas turbine engines. The new injectors offer significant aerodynamic design flexibility, excellent thermal performance, and scalability to various engine sizes. The injectors built for this project contain 15 injection points and incorporate staging to enable operation at low power conditions. Ignition and flame stability were demonstrated at ambient conditions with ignition air pressure drop as low as 0.3% and fuel-to-air ratio (FAR) as low as 0.011. Lean Blowout (LBO) occurred at FAR as low as 0.005 with air at 460 K and atmospheric pressure. A high pressure combustion testing campaign was conducted in the CE-5 test facility at NASA Glenn Research Center at pressures up to 250 psi and combustor exit temperatures up to 2,033 K (3,200 °F). The tests demonstrated estimated LTO cycle emissions that are about 30% of CAEP/6 for a reference 60,000 lbf thrust, 54.8-OPR engine. This paper presents some details of the injector design along with results from ignition, LBO and emissions testing.

Relight Envelope of a Military Gas Turbine Engine: An Experimental Study

2008 ◽  
Author(s):  
R. K. Mishra ◽  
G. Gouda ◽  
B. S. Vedaprakash
Keyword(s):  
Gas Turbine ◽  
Fuel Injection ◽  
Full Scale ◽  
Test Facility ◽  
Turbine Engines ◽  
Engine Fuel ◽  
Gas Turbine Engines ◽  
Turbine Engine ◽  
Start Up

A twin spool low bypass turbofan engine under development and its combustor in full-scale were tested independently at altitude conditions to establish the relight envelope of the engine. Demonstration of relight capability and defining its boundary are mandatory for military gas turbine engines and for single engine application in particular. The engine was first subjected to windmill to establish its windmilling characteristics. The full engine was then tested for light-off in an altitude test facility simulating windmilling conditions from 4 to 12 km altitude with flight Mach numbers from 0.2 to 1.0. The relight boundary is defined based on the successful light-off points achieved from engine tests. Similar tests were carried out on the full-scale combustion chamber in a stand-alone mode simulating altitude conditions at engine flame-out. The combustor test has defined the light-off and lean blow out limits of the at each point on the relight boundary. The information of fuel-air ratio at light-off and blow-out is very useful in setting the engine fuel schedule for altitude operation and relight. In this paper an attempt is made to highlight various tests carried out on engine and its combustor to define the relight boundary of the engine. The paper also emphasizes the experience of combustor development and associated problems in meeting the relight challenges of military engines. These problems include the necessity of higher fuel-air ratio at high altitudes, the role of additional localized fuel injection through start-up atomizers, and effect of single igniter on relight characteristics. The relight envelope demonstrated by the engine is very satisfactory to meet the first flight requirement where the flight mission generally concentrate in the zone of 0.6 to 0.8 Mach and altitude does not exceed 10 to 12 km. Combustor and atomizer modification is needed to improve relight performance and to shift the boundary to further left.

Lean Blowout Research in a Generic Gas Turbine Combustor With High Optical Access

1993 ◽  
Author(s):  
G. J. Sturgess ◽  
D. Shouse
Keyword(s):  
Experimental Data ◽  
Gas Turbine ◽  
Air Force ◽  
Operating Conditions ◽  
Test Facility ◽  
Turbine Engines ◽  
Flame Stability ◽  
Gas Turbine Engines ◽  
Gas Turbine Combustor ◽  
Lean Blowout

The U.S. Air Force is conducting a comprehensive research program aimed at improving the design and analysis capabilities for flame stability and lean blowout in the combustors of aircraft gas turbine engines. As part of this program, a simplified version of a generic gas turbine combustor is used. The intent is to provide an experimental data base against which lean blowout modeling might be evaluated and calibrated. The design features of the combustor and its instrumentation are highlighted, and the test facility is described. Lean blowout results for gaseous propane fuel are presented over a range of operating conditions at three different dome flow splits. Comparison of results with those of a simplified research combustor is also made. Lean blowout behavior is complex, so that simple phenomenological correlations of experimental data will not be general enough for use as design tools.

An Advanced Multi-Configuration Stator Well Cooling Test Facility

2010 ◽  
Author(s):  
D. D. Coren ◽  
N. R. Atkins ◽  
J. R. Turner ◽  
D. E. Eastwood ◽  
S. Davies ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Flow Rate ◽  
Coolant Flow ◽  
Coolant Flow Rate ◽  
Test Facility ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Displacement Sensors ◽  
Cooling Air ◽  
Coolant Delivery

Optimisation of cooling systems within gas turbine engines is of great interest to engine manufacturers seeking gains in performance, efficiency and component life. The effectiveness of coolant delivery is governed by complex flows within the stator wells and the interaction of main annulus and cooling air in the vicinity of the rim seals. This paper reports the development of a test facility which allows the interaction of cooling air and main gas paths to be measured at conditions representative of those found in modern gas turbine engines. The test facility features a two stage turbine with an overall pressure ratio of approximately 2.6:1. Hot air is supplied to the main annulus using a Rolls-Royce Dart compressor driven by an aero-derivative engine plant. Cooling air can be delivered to the stator wells at multiple locations and at a range of flow rates which cover bulk ingestion through to bulk egress. The facility has been designed with adaptable geometry to enable rapid changes of cooling air path configuration. The coolant delivery system allows swift and accurate changes to the flow settings such that thermal transients may be performed. Particular attention has been focused on obtaining high accuracy data, using a radio telemetry system, as well as thorough through-calibration practices. Temperature measurements can now be made on both rotating and stationary discs with a long term uncertainty in the region of 0.3 K. A gas concentration measurement system has also been developed to obtain direct measurement of re-ingestion and rim seal exchange flows. High resolution displacement sensors have been installed in order to measure hot running geometry. This paper documents the commissioning of a test facility which is unique in terms of rapid configuration changes, non-dimensional engine matching and the instrumentation density and resolution. Example data for each of the measurement systems is presented. This includes the effect of coolant flow rate on the metal temperatures within the upstream cavity of the turbine stator well, the axial displacement of the rotor assembly during a commissioning test, and the effect of coolant flow rate on mixing in the downstream cavity of the stator well.

The Requirements for a Modern Marine Gas Turbine and the Development of the WR21

1999 ◽  
Author(s):  
Brian Butler ◽  
Nigel Wright
Keyword(s):  
Gas Turbine ◽  
Development Program ◽  
Test Facility ◽  
Turbine Engines ◽  
Data Handling ◽  
Gas Turbine Engines ◽  
Royal Navy ◽  
The United Kingdom ◽  
Duty Cycles ◽  
Maritime Environment

The maritime engine test facility at the Defence Evaluation and Research Agency (DERA) Pyestock in the United Kingdom has been involved with the testing of gas turbine engines for the Royal Navy since 1952. This testing has been aimed at proving that the already developed gas turbine can operate in the maritime environment, perform the expected duty cycles, and reach the required levels of availability, reliability, and maintainability as required. To this end, the vast amount of testing at DERA has been the endurance style where the engine is run over extended periods in arduous conditions to test the life of components and with only a minimum amount of data collection. This dramatically changed in 1994 with the beginning of the WR21 Intercooled Recuperated gas turbine (WR21 ICR) development program. The paper describes the benefits of the WR21 ICR over a simple cycle gas turbine, the methodology behind the current development testing, and the evolution of the DERA Pyestock facility. In particular facility designs, instrumentation, data handling, and testing operations to meet the growing needs of the customer to develop this complex cycle gas turbine.

Tip Clearance Control Concept in Gas Turbine H.P. Compressors

2012 ◽  
Author(s):  
G. I. Ekong ◽  
C. A. Long ◽  
P. R. N. Childs
Keyword(s):  
Gas Turbine ◽  
Tip Clearance ◽  
Test Facility ◽  
Turbine Engines ◽  
Thermodynamic Efficiency ◽  
Gas Turbine Engines ◽  
Critical Importance ◽  
Operating Cycle ◽  
Expansion Time ◽  
Clearance Control

To improve the thermodynamic efficiency of aircraft engine and other gas turbine engines, higher and higher pressure ratios are desired in conjunction with more refined engine cycles. In the high pressure compressor, higher pressure ratios result in lower aspect blades and enhanced sensitivity of the engine design to radial clearance effects. The tip clearance in the axial flow compressor of modern commercial civil aero-engines is of critical importance in terms of both mechanical integrity and performance. Typically 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, and the clearance is therefore of critical importance to civil airline operators and their customers alike. A design exercise was performed and a series of conceptual solutions were developed using the theory of inventive problem solving (TRIZ) process and their potential viability in clearance control was investigated with thermal modelling. TRIZ was selected as an appropriate tool as the issue was long-standing having been the focus of previous projects, and robust design solutions were being sought. In order to validate the concepts, use was made of a test facility developed at the University of Sussex, incorporating a rotor and an inner shaft scaled down from a Rolls Royce Trent aeroengine to a ratio of 0.7:1. The mechanical design of the test facility allows the simulation of flow conditions in the HP compressor cavity equivalent to the Trent 1000 aero-engine, 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 finite element thermomechanical model has been built to validate the engine measurements. This paper describes the use of TRIZ and the development of a selected concept and the detailed evaluation for reduction and control of tip clearance in HP compressors. This was achieved through the reduction in the compressor disc heat expansion time constant by improving drum heat transfer using bleed air from the compressor core flow. This paper explores the trade-offs between clearance and efficiency and develops and explores concepts to control the compressor tip clearance throughout the engine operating cycle. The project involved modelling of potential solutions and use of experimental facilities, a rotating compressor cavity rig, in order to explore the physical principles and demonstrate proof of concept for controlling tip clearance in HP compressors of gas turbine engines.

An Advanced Multiconfiguration Stator Well Cooling Test Facility

2012 ◽  
Vol 135 (1) ◽  
Author(s):  
D. D. Coren ◽  
N. R. Atkins ◽  
J. R. Turner ◽  
D. E. Eastwood ◽  
S. Davies ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Flow Rate ◽  
Coolant Flow ◽  
Coolant Flow Rate ◽  
Test Facility ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Displacement Sensors ◽  
Cooling Air ◽  
Coolant Delivery

Optimization of cooling systems within gas turbine engines is of great interest to engine manufacturers seeking gains in performance, efficiency, and component life. The effectiveness of coolant delivery is governed by complex flows within the stator wells and the interaction of main annulus and cooling air in the vicinity of the rim seals. This paper reports on the development of a test facility which allows the interaction of cooling air and main gas paths to be measured at conditions representative of those found in modern gas turbine engines. The test facility features a two stage turbine with an overall pressure ratio of approximately 2.6:1. Hot air is supplied to the main annulus using a Rolls-Royce PLC Dart compressor driven by an aero-derivative engine plant. Cooling air can be delivered to the stator wells at multiple locations and at a range of flow rates which cover bulk ingestion through to bulk egress. The facility has been designed with adaptable geometry to enable rapid changes of cooling air path configuration. The coolant delivery system allows swift and accurate changes to the flow settings such that thermal transients may be performed. Particular attention has been focused on obtaining high accuracy data, using a radio telemetry system, as well as thorough through-calibration practices. Temperature measurements can now be made on both rotating and stationary disks with a long term uncertainty in the region of 0.3 K. A gas concentration measurement system has also been developed to obtain direct measurement of re-ingestion and rim seal exchange flows. High resolution displacement sensors have been installed in order to measure hot running geometry. This paper documents the commissioning of a test facility which is unique in terms of rapid configuration changes, nondimensional engine matching, and the instrumentation density and resolution. Example data for each of the measurement systems are presented. This includes the effect of coolant flow rate on the metal temperatures within the upstream cavity of the turbine stator well, the axial displacement of the rotor assembly during a commissioning test, and the effect of coolant flow rate on mixing in the downstream cavity of the stator well.

Lean Blowout Research in a Generic Gas Turbine Combustor With High Optical Access

1997 ◽  
Vol 119 (1) ◽  
pp. 108-118 ◽  
Author(s):  
G. J. Sturgess ◽  
D. Shouse
Keyword(s):  
Experimental Data ◽  
Gas Turbine ◽  
Air Force ◽  
Operating Conditions ◽  
Test Facility ◽  
Turbine Engines ◽  
Flame Stability ◽  
Gas Turbine Engines ◽  
Gas Turbine Combustor ◽  
Lean Blowout

The U. S. Air Force is conducting a comprehensive research program aimed at improving the design and analysis capabilities for flame stability and lean blowout in the combustors of aircraft gas turbine engines. As part of this program, a simplified version of a generic gas turbine combustor is used. The intent is to provide an experimental data base against which lean blowout modeling might be evaluated and calibrated. The design features of the combustor and its instrumentation are highlighted, and the test facility is described. Lean blowout results for gaseous propane fuel are presented over a range of operating conditions at three different dome flow splits. Comparison of results with those of a simplified research combustor is also made. Lean blowout behavior is complex, so that simple phenomenological correlations of experimental data will not be general enough for use as design tools.

Sign in / Sign up

Export Citation Format

Share Document