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.

A Gas Turbine Blade Thermal/Structural Program With Linked Flow-Solid Modeling Capability

1994 ◽  
Author(s):  
S. M. Wan ◽  
T. C. T. Lam ◽  
J. M. Allen ◽  
T. H. McCloskey
Keyword(s):  
Steady State ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Gas Flow ◽  
Thermal Stresses ◽  
Turbine Blades ◽  
Solid Modeling ◽  
Mechanical Analysis ◽  
Operating Conditions ◽  
Gas Turbine Blade

A time-marching approach is adopted in developing a thermal/structural program with linked flow-solid modeling capability. The Blade Life Analysis & Design Evaluation for Combustion Turbines (BLADE-CT) program analyzes gas turbine blade thermal-mechanical stress and natural frequencies under the boundary conditions which result from the gas flow and the cooling/barrier flow within a given turbine stage. Using the finite element method, the blade temperatures obtained from transient/steady-state thermal solutions can be utilized to compute thermal stresses and dynamic stresses under operating conditions for assessing thermal-mechanical fatigue damage in combustion turbine blades. A customized and automated mesh generation routine is developed to model cooled (spanwise multihole configurations) and solid gas turbine blades. By coupling the NASA flow programs, PCPANEL (potential flow), STAN5 (heat transfer boundary layer), and CPF (coolant passage flow) as part of an automated flow-structural analysis approach, a more efficient and accurate thermal and thermal stress calculation can be achieved. The calculated blade temperatures can be also applied for the frequency analysis to account for temperature effects. The coupled fluid-structure interaction program approach for thermal-mechanical analysis and an example of a spanwise cooled blade steady state analysis are presented.

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.

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.

Development of the Ceramic Gas Turbine Engine System (CGT301)

1999 ◽  
Author(s):  
Takeshi Sakida ◽  
Shinya Tanaka ◽  
Takao Mikami ◽  
Masashi Tatsuzawa ◽  
Tomoki Taoka
Keyword(s):  
Gas Turbine ◽  
Turbine Blades ◽  
Gas Turbine Engine ◽  
Turbine Engine ◽  
The Future ◽  
Engine System ◽  
Primary Type ◽  
Three Phases

The CGT301 ceramic gas turbine has been developed under a contract from NEDO as a part of the New Sunshine Program of MITI since 1988 to 1998. The CGT301 is a recuperated, single-shaft ceramic gas turbine. Ceramic parts are used in the hot section of the engine, such as turbine blades, nozzle vanes, combustion liners and so on. As a primary feature of this turbine, the rotors are composed of ceramic blades inserted into metallic disks (“hybrid rotor”) for the future applicability to the large gas turbine. The R & D program consists of three phases, the model metal gas turbine, the primary type ceramic gas turbine and the pilot ceramic gas turbine. The pilot ceramic gas turbine showed etable operation at TIT of 1,350°C. This paper presents the progress in the development of the pilot ceramic gas turbine of CGT301.

Fabrication and Performance of a Pin Fin Micro Heat Exchanger

2004 ◽  
Vol 126 (3) ◽  
pp. 434-444 ◽  
Author(s):  
Christophe Marques ◽  
Kevin W. Kelly
Keyword(s):  
Heat Exchanger ◽  
Gas Turbine ◽  
Heat Exchangers ◽  
Turbine Blades ◽  
High Heat ◽  
Gas Turbine Engine ◽  
High Heat Flux ◽  
Turbine Engine ◽  
Pin Fin ◽  
Micro Heat Exchanger

Nickel micro pin fin heat exchangers can be electroplated directly onto planar or non-planar metal surfaces using a derivative of the LIGA micromachining process. These heat exchangers offer the potential to more effectively control the temperature of surfaces in high heat flux applications. Of particular interest is the temperature control of gas turbine engine components. The components in the gas turbine engine that require efficient, improved cooling schemes include the gas turbine blades, the stator vanes, the turbine disk, and the combustor liner. Efficient heating of component surfaces may also be required (i.e., surfaces near the compressor inlet to prevent deicing). In all cases, correlations providing the Nusselt number and the friction factor are needed for such micro pin fin heat exchangers. Heat transfer and pressure loss experimental results are reported for a flat parallel plate pin fin micro heat exchanger with a staggered pin fin array, with height-to-diameter ratios of 1.0, with spacing-to-diameter ratios of 2.5 and for Reynolds numbers (based on the hydraulic diameter of the channel) from 4000 to 20,000. The results are compared to studies of larger scale, but geometrically similar, pin fin heat exchangers. To motivate further research, an analytic model is described which uses the empirical results from the pin fin heat exchanger experiments to predict a cooling effectiveness exceeding 0.82 in a gas turbine blade cooling application. As a final point, the feasibility of fabricating a relatively complex micro heat exchanger on a simple airfoil (a cylinder) is demonstrated.

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