scholarly journals Salt-Water Problems in Marine Gas Turbines

1965 ◽  
Author(s):  
Eugene P. Weinert ◽  
Gilbert A. Carlton
Keyword(s):  
Solid State ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Salt Water ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Sea Salt ◽  
Turbine Engine ◽  
Compressor Performance ◽  
Water Problems

The gas-turbine engine in naval service is subjected to severe operating conditions. By far the most consistent troublemaker is sea salt. In either liquid or solid state, sea salt causes problems in corrosion of hot and cold surfaces, fouling of fuel systems, and deterioration of compressor performance. Such problems are reviewed and solutions discussed.

Exploration of Compact Combustors for Reheat Cycle Aero Engine Applications

2006 ◽  
Author(s):  
J. Zelina ◽  
D. T. Shouse ◽  
J. S. Stutrud ◽  
G. J. Sturgess ◽  
W. M. Roquemore
Keyword(s):  
Gas Turbine ◽  
Temperature Rise ◽  
Gas Turbines ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Combustion System ◽  
Turbine Engine ◽  
Lean Blowout ◽  
Power Extraction ◽  
Wide Range

An aero gas turbine engine has been proposed that uses a near-constant-temperature (NCT) cycle and an Inter-Turbine Burner (ITB) to provide large amounts of power extraction from the low-pressure turbine. This level of energy is achieved with a modest temperature rise across the ITB. The additional energy can be used to power a large geared fan for an ultra-high bypass ratio transport aircraft, or to drive an alternator for large amounts of electrical power extraction. Conventional gas turbines engines cannot drive ultra-large diameter fans without causing excessively high turbine temperatures, and cannot meet high power extraction demands without a loss of engine thrust. Reducing the size of the combustion system is key to make use of a NCT gas turbine cycle. Ultra-compact combustor (UCC) concepts are being explored experimentally. These systems use high swirl in a circumferential cavity about the engine centerline to enhance reaction rates via high cavity g-loading on the order of 3000 g’s. Any increase in reaction rate can be exploited to reduce combustor volume. The UCC design integrates compressor and turbine features which will enable a shorter and potentially less complex gas turbine engine. This paper will present experimental data of the Ultra-Compact Combustor (UCC) performance in vitiated flow. Vitiation levels were varied from 12–20% oxygen levels to simulate exhaust from the high pressure turbine (HPT). Experimental results from the ITB at atmospheric pressure indicate that the combustion system operates at 97–99% combustion efficiency over a wide range of operating conditions burning JP-8 +100 fuel. Flame lengths were extremely short, at about 50% of those seen in conventional systems. A wide range of operation is possible with lean blowout fuel-air ratio limits at 25–50% below the value of current systems. These results are significant because the ITB only requires a small (300°F) temperature rise for optimal power extraction, leading to operation of the ITB at near-lean-blowout limits of conventional combustor designs. This data lays the foundation for the design space required for future engine designs.

scholarly journals Prediction and Analysis of Impact of Thermal Barrier Coating Oxidation on Gas Turbine Creep Life

2016 ◽  
Vol 138 (12) ◽  
Author(s):  
E. A. Ogiriki ◽  
Y. G. Li ◽  
Th. Nikolaidis
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Thermal Barrier ◽  
Creep Life ◽  
Premature Failure ◽  
Turbine Engine ◽  
Oxidation Rates ◽  
The Impact

Thermal barrier coatings (TBCs) have been widely used in the power generation industry to protect turbine blades from damage in hostile operating environment. This allows either a high turbine entry temperature (TET) to be employed or a low percentage of cooling air to be used, both of which will improve the performance and efficiency of gas turbine engines. However, with continuous increases in TET aimed at improving the performance and efficiency of gas turbines, TBCs have become more susceptible to oxidation. Such oxidation has been largely responsible for the premature failure of most TBCs. Nevertheless, existing creep life prediction models that give adequate considerations to the effects of TBC oxidation on creep life are rare. The implication is that the creep life of gas turbines may be estimated more accurately if TBC oxidation is considered. In this paper, a performance-based integrated creep life model has been introduced with the capability of assessing the impact of TBC oxidation on the creep life and performance of gas turbines. The model comprises of a thermal, stress, oxidation, performance, and life estimation models. High pressure turbine (HPT) blades are selected as the life limiting component of gas turbines. Therefore, the integrated model was employed to investigate the effect of several operating conditions on the HPT blades of a model gas turbine engine using a creep factor (CF) approach. The results show that different operating conditions can significantly affect the oxidation rates of TBCs which in turn affect the creep life of HPT blades. For instance, TBC oxidation can speed up the overall life usage of a gas turbine engine from 4.22% to 6.35% within a one-year operation. It is the objective of this research that the developed method may assist gas turbine users in selecting the best mission profile that will minimize maintenance and operating costs while giving the best engine availability.

Prediction and Analysis of Impact of TBC Oxidation on Gas Turbine Creep Life

2015 ◽  
Author(s):  
E. A. Ogiriki ◽  
Y. G. Li ◽  
Th. Nikolaidis
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Prediction Models ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Creep Life ◽  
Premature Failure ◽  
Turbine Engine ◽  
Oxidation Rates ◽  
The Impact

Thermal Barrier Coatings (TBC) have been widely used in the power generation industry to protect turbine blades from damage in hostile operating environment. This allows either a high Turbine Entry temperature (TET) to be employed or a low percentage of cooling air to be used, both of which will improve the performance and efficiency of gas turbine engines. However, with continuous increases in turbine entry temperature aimed at improving the performance and efficiency of gas turbines, TBCs have become more susceptible to oxidation. Such oxidation has been largely responsible for the premature failure of most TBCs. Nevertheless, existing creep life prediction models that give adequate considerations to the effects of TBC oxidation on creep life are rare. The implication is that the creep life of gas turbines may be estimated more accurately if TBC oxidation is considered. In this paper, a performance-based integrated creep life model has been introduced with the capability of assessing the impact of TBC oxidation on the creep life and performance of gas turbines. The model comprises of a thermal, stress, oxidation, performance, and life estimation models. High Pressure Turbine (HPT) blades are selected as the life limiting component of the gas turbine. Therefore the integrated model was employed to investigate the effect of several operating conditions on the HPT blades of a model gas turbine engine using a Creep Factor approach. The results show that different operating conditions can significantly affect the oxidation rates of TBCs which in turn affect the creep life of HPT blades. For instance, TBC oxidation can speed up the overall life usage of a gas turbine engine from 4.22% to 6.35% within one year operation. It is the objective of this research that the developed method may assist gas turbine users in selecting the best mission profile that will minimize maintenance and operating costs while giving the best engine availability.

Optimal State-Space Control of a Gas Turbine Engine

1992 ◽  
Vol 114 (4) ◽  
pp. 763-767 ◽  
Author(s):  
J. W. Watts ◽  
T. E. Dwan ◽  
C. G. Brockus
Keyword(s):  
State Space ◽  
Gas Turbine ◽  
State Space Model ◽  
Gas Turbine Engine ◽  
The State ◽  
Operating Conditions ◽  
Turbine Engine ◽  
Analog Control ◽  
Linear State ◽  
High Level

An analog fuel control for a gas turbine engine was compared with several state-space derived fuel controls. A single-spool, simple cycle gas turbine engine was modeled using ACSL (high level simulation language based on FORTRAN). The model included an analog fuel control representative of existing commercial fuel controls. The ACSL model was stripped of nonessential states to produce an eight-state linear state-space model of the engine. The A, B, and C matrices, derived from rated operating conditions, were used to obtain feedback control gains by the following methods: (1) state feedback; (2) LQR theory; (3) Bellman method; and (4) polygonal search. An off-load transient followed by an on-load transient was run for each of these fuel controls. The transient curves obtained were used to compare the state-space fuel controls with the analog fuel control. The state-space fuel controls did better than the analog control.

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 Comprehensive report on Materials for Gas Turbine Engine Components

2019 ◽  
Vol 9 (2S2) ◽  
pp. 155-158
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Thermal Stresses ◽  
Prolonged Exposure ◽  
Raw Materials ◽  
Turbine Blades ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Turbine Engine ◽  
Engine Component

In the past three decades, it is very challenging for the researchers to design and development a best gas turbine engine component. Engine component has to face different operating conditions at different working environments. Nickel based superalloys are the best material to design turbine components. Inconel 718, Inconel 617, Hastelloy, Monel and Udimet are the common material used for turbine components. Directional solidification is one of the conventional casting routes followed to develop turbine blades. It is also reported that the raw materials are heat treated / age hardened to enrich the desired properties of the material implementation. Accordingly they are highly susceptible to mechanical and thermal stresses while operating. The hot section of the turbine components will experience repeated thermal stress. The halides in the combination of sulfur, chlorides and vanadate are deposited as molten salt on the surface of the turbine blade. On prolonged exposure the surface of the turbine blade starts to peel as an oxide scale. Microscopic images are the supportive results to compare the surface morphology after complete oxidation / corrosion studies. The spectroscopic results are useful to identify the elemental analysis over oxides formed. The predominant oxides observed are NiO, Cr2O3, Fe2O3 and NiCr2O4. These oxides are vulnerable on prolonged exposure and according to PB ratio the passivation are very less. In recent research, the invention on nickel based superalloys turbine blades produced through other advanced manufacturing process is also compared. A summary was made through comparing the conventional material and advanced materials performance of turbine blade material for high temperature performance.

scholarly journals Comparison of Coriolis and Turbine Type Flow Meters for Fuel Measurement in Gas Turbine Testing

1993 ◽  
Author(s):  
J. D. MacLeod ◽  
W. Grabe
Keyword(s):  
Steady State ◽  
Gas Turbine ◽  
Mass Flow ◽  
Gas Turbines ◽  
Gas Turbine Engine ◽  
Turbine Engine ◽  
Turbine Flowmeter ◽  
Turbine Performance ◽  
Coriolis Flowmeter ◽  
Project Objectives

The Machinery and Engine Technology (MET) Program of the National Research Council of Canada (NRCC) has established a program for the evaluation of sensors to measure gas turbine engine performance accurately. The precise measurement of fuel flow is an essential part of steady-state gas turbine performance assessment. Prompted by an international engine testing and information exchange program, and a mandate to improve all aspects of gas turbine performance evaluation, the MET Laboratory has critically examined two types of fuel flowmeters, Coriolis and turbine. The two flowmeter types are different in that the Coriolis flowmeter measures mass flow directly, while the turbine flowmeter measures volumetric flow, which must be converted to mass flow for conventional performance analysis. The direct measurement of mass flow, using a Coriolis flowmeter, has many advantages in field testing of gas turbines, because it reduces the risk of errors resulting from the conversion process. Turbine flowmeters, on the other hand, have been regarded as an industry standard because they are compact, rugged, reliable, and relatively inexpensive. This paper describes the project objectives, the experimental installation, and the results of the comparison of the Coriolis and turbine type flowmeters in steady-state performance testing. Discussed are variations between the two types of flowmeters due to fuel characteristics, fuel handling equipment, acoustic and vibration interference and installation effects. Also included in this paper are estimations of measurement uncertainties for both types of flowmeters. Results indicate that the agreement between Coriolis and turbine type flowmeters is good over the entire steady-state operating range of a typical gas turbine engine. In some cases the repeatability of the Coriolis flowmeter is better than the manufacturers specification. Even a significant variation in fuel density (10%), and viscosity (300%), did not appear to compromise the ability of the Coriolis flowmeter to match the performance of the turbine flowmeter.

Method for calculating the thrust and specific thrust of a double-flow gas turbine engine based on the results of tests at the laboratory unit «Flame»

2021 ◽  
pp. 58-64
Author(s):  
S.M. Sergeev ◽  
◽  
V.A. Kudriashov ◽  
N.V. Petrukhin ◽  
◽  
...  
Keyword(s):  
Gas Turbine ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Similarity Criteria ◽  
Fuel Quality ◽  
Jet Engines ◽  
Laboratory Unit ◽  
Turbine Engine ◽  
Specific Thrust ◽  
The Cost

The main technical characteristics of jet engines depend on the fuel quality: thrust and fuel consumption. As a rule, the comparative assessment of real engines is carried by specific values. Specific thrust is one of the most important parameters of the gas turbine engine (GTE). The larger it is, the smaller the required air flow rate through the engine at a given thrust and therefore its dimensions and mass. To date, a system for evaluating the performance properties of fuels based on qualification methods has been created. However, these methods do not allow calculating the thrust and specific thrust of the engine and potentially assessing the effect of fuels on these characteristics. Therefore, the issues of efficient use of fuels for GTE are solved almost exclusively on the basis of tests at testing units with full-scale engines, which are carried out repeatedly, which leads to a significant increase in the cost of testing. The article proposes a method for calculating the thrust and specific thrust of a double-flow gas turbine engine according to the results of tests at a constant volume laboratory unit of bypass type “Flame”. The method is based on modeling the engine operating conditions using the similarity criteria of the bench reactor and the real engine and allows reducing significantly the material and time costs for testing. The experimental of the combustion characteristics of hydrocarbon fuels and the rated values of their thrust and specific thrust for a double-flow gas turbine engine are presented.

Cycle Design Exploration Using Multi-Design Point Approach

2012 ◽  
Author(s):  
Jeffrey Schutte ◽  
Jimmy Tai ◽  
Jonathan Sands ◽  
Dimitri Mavris
Keyword(s):  
Gas Turbine ◽  
Design Space ◽  
Traditional Approach ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Feasible Region ◽  
Turbine Engine ◽  
Design Point ◽  
Performance Requirements ◽  
Design Cycle

The focus of this study is to compare the aerothermodynamic cycle design space of a gas turbine engine generated using two on-design approaches. The traditional approach uses a single design point (SDP) for on-design cycle analysis, where off-design cycle analysis must be performed at other operating conditions of interest. A multi-design point (MDP) method performs on-design cycle analysis at all operating conditions where performance requirements are specified. Effects on the topography of the cycle design space as well as the feasibility of the space are examined. The impacts that performance requirements and cycle assumptions have on the bounds and topography of the feasible space are investigated. The deficiencies of a SDP method in determining an optimum gas turbine engine will be shown for a given set of requirements. Analysis will demonstrate that the MDP method, unlike the SDP method, always obtains a properly sized engine for a set of given requirements and cycle design variables, resulting in an increased feasible region of the aerothermodynamic cycle design space from which the optimum performance engine can be obtained.

Compensation of Intake Induced Flow Non-Uniformities With Compressor Blade Lean Angle Changes

2009 ◽  
Author(s):  
Ioannis Templalexis ◽  
Vassilios Pachidis ◽  
Petros Kotsiopoulos
Keyword(s):  
Gas Turbine ◽  
Flow Simulation ◽  
Gas Turbine Engine ◽  
Compressor Blade ◽  
Design Stage ◽  
Computational Time ◽  
Turbine Engine ◽  
Lean Angle ◽  
Compressor Performance ◽  
Induced Flow

The compression system has traditionally drawn most of the attention concerning the gas turbine engine performance assessment and design procedure. It is the most vulnerable component to flow fluctuations within a gas turbine engine. In particular this study focuses on performance deviations, between an installed and an uninstalled compressor. Test results acquired from a test bed installation will differ from these recorded when the compressor operates as an integral part of the engine. The upstream duct, whether an intake or an interstage duct, will affect the flow field pattern ingested into the compressor. The case studies presented into this work aim to mostly qualify the effect of boundary layer growth along the upstream duct walls, upon compressor performance. Additionally, compressor performance response on blade lean angle variation is being addressed, with the aim of acquiring an understanding as to how compressor blade lean angle changes interact with intake induced flow non uniformities. Such studies are usually conducted during the preliminary design stage, before the compressor is built. Consequently, experimental performance investigation is excluded at this stage of development. Computer aided simulation techniques are between the few if not the only option for compressor performance prediction. Given the fact that many such design parameters need to be assessed under the time pressure exerted by the tight compressor development program, the compressor flow simulation technique used needs to provide reliable results while consuming the least possible computational time. Such a low computational time compressor flow simulation method, among others, is the two dimensional (2D) streamline curvature (SLC) method, being applied within the frame of reference of the current study. The paper is introduced by a brief discussion on SLC method that was proposed more than 50 years ago. Then a reference is made to the Radial Equilibrium Equation (REE) which is the mathematical basis of the code, commenting on the assumptions that were undertaken. Subsequently the influence of the intake presence on the compressor inlet radial flow distribution is being addressed, with the aim of adjusting compressor blade inlet lean angle, in order to minimize compressor performance deterioration. Finally the paper is concluded with a discussion of the results.

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