scholarly journals Multi-Purpose Fuel: Problems That We Face

1969 ◽  
Author(s):  
John S. Siemietkowski
Keyword(s):  
Turbine Blades ◽  
Engine Performance ◽  
Problem Area ◽  
Turbine Engine ◽  
Short Period ◽  
Marine Diesel ◽  
Fuel Filter ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Cycle Endurance

A Pratt and Whitney FT4A Marine Gas Turbine Engine rated at 25,000 HP for a 100°F inlet day, was tested at the Naval Ship Engineering Center Philadelphia Division for a total of 201 hours, 15 minutes. The engine was subjected to an initial 30 hour “coking” run, conducted at 10,000 HP, 2380 rpm, to determine adverse effects on the engine under simulated destroyer type operation. Following the 30 hour coking run, the engine was subjected to a 150 hour cycle endurance operation. Salt was admitted to the inlet air. A combustion section inspection was performed at the end of the 30 hour coking run. No detrimental effects were noted at that time. An overall combustion section inspection was performed at the end of the test. A fuel manifold and nozzle spray check was performed with both acceptable for further use. First stage turbine blades showed some degree of sulfidation, while the nozzle guide vanes showed evidence of coating loss and partial penetration into the base metal (on only the uncoated vanes). The major problem area during the test was the failure of the coalescer fuel filter to function properly with Multi-Purpose fuel. Due to the higher pour point (with attendant “wax” precipitation) of the fuel in comparison with normal Marine Diesel (MIL-F-16884), plugging of the coalescer filter elements occurred in a very short period of time. Engine performance over the entire test was satisfactory approximating that of previous FT4A testing.

Experimental Analysis of an Ultra Compact Combustor Powered Turbine Engine

2019 ◽  
Author(s):  
Brian T. Bohan ◽  
Marc D. Polanka
Keyword(s):  
Engine Performance ◽  
Combustion Products ◽  
Combustion Efficiency ◽  
Small Scale ◽  
Engine Operation ◽  
Turbine Engine ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Length Reduction ◽  
Outside Diameter

Abstract The innovative Ultra Compact Combustor (UCC) is an alternative to traditional turbine engine combustors and has been shown to reduce the combustor volume and offer potential improvements in combustion efficiency. Prior UCC configurations featured a circumferential combustion cavity positioned around the outside diameter (OD) of the engine. This configuration would be difficult to implement in a vehicle with a small, fixed diameter and had difficulty migrating the hot combustion products at the OD radially inward across an axial core flow to present a uniform temperature distribution to the first turbine stage. The present study experimentally tested a new UCC configuration that featured a circumferential cavity that exhausted axially into a dilution zone positioned just upstream of the nozzle guide vanes. The combustor was sized as a replacement burner for the JetCat P90 RXi small-scale turbine engine and fit inside the engine casing. This combustor configuration achieved a 33% length reduction compared to the stock JetCat combustor and achieved comparable engine performance across a limited operating range. Self-sustaining engine operation was achieved with a rotating compressor and turbine making this study the first to achieve operation of a UCC powered turbine engine.

Experimental Analysis of an Ultra-Compact Combustor Powered Turbine Engine

2020 ◽  
Vol 142 (5) ◽  
Author(s):  
Brian T. Bohan ◽  
Marc D. Polanka
Keyword(s):  
Experimental Testing ◽  
Engine Performance ◽  
Combustion Products ◽  
Combustion Efficiency ◽  
Small Scale ◽  
Engine Operation ◽  
Turbine Engine ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Length Reduction

Abstract The innovative ultra-compact combustor (UCC) is an alternative to traditional turbine engine combustors and has been shown to reduce the combustor length and offer potential improvements in combustion efficiency. Prior UCC configurations featured a circumferential combustion cavity positioned around the outside diameter (OD) of the engine. This configuration would be difficult to implement in a vehicle with a small, fixed diameter and had difficulty migrating the hot combustion products at the OD radially inward across an axial core flow to present a uniform temperature distribution to the first turbine stage. This study draws from preliminary computational analysis which enabled experimental testing of a new UCC configuration that featured a smaller diameter circumferential cavity that exhausted axially into a dilution zone positioned just upstream of the nozzle guide vanes. The combustor was sized as a replacement burner for the JetCat P90 RXi small-scale turbine engine and fit inside the engine casing. This combustor configuration achieved a 33% length reduction compared to the stock JetCat combustor and achieved comparable engine performance across a limited operating range. Self-sustained engine operation was achieved with a rotating compressor and turbine making this study the first to achieve operation of a UCC-powered turbine engine.

Flow Field Simulations of a Gas Turbine Combustor

2002 ◽  
Vol 124 (3) ◽  
pp. 508-516 ◽  
Author(s):  
M. D. Barringer ◽  
O. T. Richard ◽  
J. P. Walter ◽  
S. M. Stitzel ◽  
K. A. Thole
Keyword(s):  
Flow Field ◽  
Gas Turbine ◽  
Large Scale ◽  
Guide Vane ◽  
Wall Region ◽  
Turbine Engine ◽  
Tunnel Section ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Combustor Liner

The flow field exiting the combustor in a gas turbine engine is quite complex considering the presence of large dilution jets and complicated cooling schemes for the combustor liner. For the most part, however, there has been a disconnect between the combustor and turbine when simulating the flow field that enters the nozzle guide vanes. To determine the effects of a representative combustor flow field on the nozzle guide vane, a large-scale wind tunnel section has been developed to simulate the flow conditions of a prototypical combustor. This paper presents experimental results of a combustor simulation with no downstream turbine section as a baseline for comparison to the case with a turbine vane. Results indicate that the dilution jets generate turbulence levels of 15–18% at the exit of the combustor with a length scale that closely matches that of the dilution hole diameter. The total pressure exiting the combustor in the near-wall region neither resembles a turbulent boundary layer nor is it completely uniform putting both of these commonly made assumptions into question.

Validation and Comparison of Strip Seal Designs for Gas Turbine Engine Nozzle Guide Vanes

2008 ◽  
Author(s):  
Arash Farahani ◽  
Peter Childs
Keyword(s):  
Gas Turbine ◽  
Guide Vane ◽  
Gas Turbine Engine ◽  
Experimental Comparison ◽  
Cfd Modelling ◽  
Guide Vanes ◽  
Turbine Engine ◽  
Seal Design ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Strip seals are commonly used to prevent or limit leakage flows between nozzle guide vanes (NGV) and other gas turbine engine components that are assembled from individual segments. Leakage flow across, for example, a nozzle guide vane platform, leads to increased demands on the gas turbine engine internal flow system and a rise in specific fuel consumption (SFC). Careful attention to the flow characteristics of strip seals is therefore necessary. The very tight tolerances associated with strip seals provides a particular challenge to their characterisation. This paper reports the validation of CFD modelling for the case of a strip seal under very carefully controlled conditions. In addition, experimental comparison of three types of strip seal design, straight, arcuate, and cloth, is presented. These seals are typical of those used by competing manufacturers of gas turbine engines. The results show that the straight seal provides the best flow sealing performance for the controlled configuration tested, although each design has its specific merits for a particular application.

The Evolution of Thermal Barrier Coatings in Gas Turbine Engine Applications

1994 ◽  
Vol 116 (1) ◽  
pp. 250-257 ◽  
Author(s):  
S. M. Meier ◽  
D. K. Gupta
Keyword(s):  
Thermal Barrier Coatings ◽  
Turbine Blades ◽  
Engine Performance ◽  
Design System ◽  
Thermal Barrier ◽  
Barrier Coatings ◽  
Turbine Engine ◽  
Component Design ◽  
Integral Element ◽  
Plasma Sprayed

Thermal barrier coatings (TBCs) have been used for almost three decades to extend the life of combustors and augmentors and, more recently, stationary turbine components. Plasma-sprayed yttria-stabilized zirconia TBC currently is bill-of-material on many commercial jet engine parts. A more durable electron beam-physical vapor deposited (EB-PVD) ceramic coating recently has been developed for more demanding rotating as well as stationary turbine components. This ceramic EB-PVD is bill-of-material on turbine blades and vanes in current high thrust engine models and is being considered for newer developmental engines as well. To take maximum advantage of potential TBC benefits, the thermal effect of the TBC ceramic layer must become an integral element of the hot section component design system. To do this with acceptable reliability requires a suitable analytical life prediction model calibrated to engine experience. The latest efforts in thermal barrier coatings are directed toward correlating such models to measured engine performance.

Prediction of Ingestion Through Turbine Rim Seals—Part I: Rotationally Induced Ingress

2010 ◽  
Vol 133 (3) ◽  
Author(s):  
J. Michael Owen
Keyword(s):  
Flow Rate ◽  
Turbine Blades ◽  
External Pressure ◽  
Swirling Flows ◽  
Radial Gradient ◽  
Mainstream Flow ◽  
Predicted Values ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Good Agreement

The mainstream flow past the stationary nozzle guide vanes and rotating turbine blades in a gas turbine creates an unsteady nonaxisymmetric variation in pressure in the annulus, radially outward of the rim seal. The ingress and egress occur through those parts of the seal clearance where the external pressure is higher and lower, respectively, than that in the wheel-space; this nonaxisymmetric type of ingestion is referred to here as externally induced (EI) ingress. Another cause of ingress is that the rotating air inside the wheel-space creates a radial gradient of pressure so that the pressure inside the wheel-space can be less than that outside; this creates rotationally induced (RI) ingress, which—unlike EI ingress—can occur, even if the flow in the annulus is axisymmetric. Although the EI ingress is usually dominant in a turbine, there are conditions under which both EI and RI ingress are significant, these cases are referred to as combined ingress. In Part I of this two-part paper, the so-called orifice equations are derived for compressible and incompressible swirling flows, and the incompressible equations are solved analytically for the RI ingress. The resulting algebraic expressions show how the nondimensional ingress and egress vary with Θ0, which is the ratio of the flow rate of sealing air to the flow rate necessary to prevent ingress. It is shown that ε, the sealing effectiveness, depends principally on Θ0, and the predicted values of ε are in mainly in good agreement with the available experimental data.

Study over thermal state of gas turbine engine metal-ceramic rotor blades and nozzle guide vanes under thermal shock and thermal-cyclic loading conditions

2004 ◽  
Vol 13 (2) ◽  
pp. 163-166
Author(s):  
A. V. Soudarev ◽  
A. A. Souryaninov ◽  
V. Ya. Podgorets ◽  
V. V. Grishaev ◽  
V.Yu Tikhoplav ◽  
...  
Keyword(s):  
Cyclic Loading ◽  
Thermal Shock ◽  
Gas Turbine ◽  
Thermal State ◽  
Rotor Blades ◽  
Guide Vanes ◽  
Turbine Engine ◽  
Metal Ceramic ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Experimental and Computational Comparisons of Fan-Shaped Film-Cooling on a Turbine Vane Surface

2005 ◽  
Author(s):  
W. Colban ◽  
K. A. Thole ◽  
M. Haendler
Keyword(s):  
Turbulence Model ◽  
Film Cooling ◽  
Turbulence Models ◽  
Complex Flow ◽  
Ir Camera ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Film Cooling Holes

The flow exiting the combustor in a gas turbine engine is considerably hotter than the melting temperature of the turbine section components, of which the turbine nozzle guide vanes see the hottest gas temperatures. One method used to cool the vanes is to use rows of film-cooling holes to inject bleed air that is lower in temperature through an array of discrete holes onto the vane surface. The purpose of this study was to evaluate the row-by-row interaction of fan-shaped holes as compared to the performance of a single row of fan-shaped holes in the same locations. This study presents adiabatic film-cooling effectiveness measurements from a scaled-up, two-passage vane cascade. High resolution film-cooling measurements were made with an infrared (IR) camera at a number of engine representative flow conditions. Computational fluid dynamics (CFD) predictions were also made to evaluate the performance of some of the current turbulence models in predicting a complex flow such as turbine film-cooling. The RNG k-ε turbulence model gave a closer prediction of the overall level of film-effectiveness, while the v2-f turbulence model gave a more accurate representation of the flow physics seen in the experiments.

Prediction of Ingestion Through Turbine Rim Seals—Part 1: Rotationally-Induced Ingress

2009 ◽  
Author(s):  
J. Michael Owen
Keyword(s):  
Flow Rate ◽  
Swirling Flow ◽  
Turbine Blades ◽  
External Pressure ◽  
Radial Gradient ◽  
Mainstream Flow ◽  
Predicted Values ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Good Agreement

The mainstream flow past the stationary nozzle guide vanes and rotating turbine blades in a gas turbine creates an unsteady non-axisymmetric variation of pressure in the annulus radially outward of the rim seal. Ingress and egress occur through those parts of the seal clearance where the external pressure is higher and lower, respectively, than that in the wheel-space; this non-axisymmetric type of ingestion is referred to here as externally-induced (EI) ingress. Another cause of ingress is that the rotating air inside the wheel-space creates a radial gradient of pressure so that the pressure inside the wheel-space can be less than that outside; this creates rotationally-induced (RI) ingress, which — unlike EI ingress — can occur even if the flow in the annulus is axisymmetric. Although EI ingress is usually dominant in a turbine, there are conditions under which both EI and RI ingress are significant: these cases are referred to as combined ingress. In Part 1 of this two-part paper, the so-called orifice equations are derived for compressible and incompressible swirling flow, and the incompressible equations are solved analytically for RI ingress. The resulting algebraic expressions show how the nondimensional ingress and egress vary with Θ0, the ratio of the flow rate of sealing air to the flow rate necessary to prevent ingress. It is shown that ε, the sealing effectiveness, depends principally on Θ0, and the predicted values of ε are in mainly good agreement with available experimental data. Part 2 (ASME GT2009-59122) concentrates on the solution and validation of the orifice equations for EI and combined ingress.

The Evolution of Thermal Barrier Coatings in Gas Turbine Engine Applications

1992 ◽  
Author(s):  
Susan Manning Meier ◽  
Dinesh K. Gupta
Keyword(s):  
Thermal Barrier Coatings ◽  
Turbine Blades ◽  
Engine Performance ◽  
Design System ◽  
Thermal Barrier ◽  
Barrier Coatings ◽  
Turbine Engine ◽  
Component Design ◽  
Integral Element ◽  
Plasma Sprayed

Thermal barrier coatings (TBCs) have been used for almost three decades to extend the life of combustors and augmentors and, more recently, stationary turbine components. Plasma sprayed yttria stabilized zirconia TBC currently is bill-of-material on many commercial jet engine parts. A more durable electron beam-physical vapor deposited (EB-PVD) ceramic coating recently has been developed for more demanding rotating as well as stationary turbine components. This ceramic EB-PVD is bill-of-material on turbine blades and vanes in current high thrust engine models and is being considered for newer developmental engines as well. To take maximum advantage of potential TBC benefits, the thermal effect of the TBC ceramic layer must become an integral element of the hot section component design system. To do this with acceptable reliability requires a suitable analytical life prediction model calibrated to engine experience. The latest efforts in thermal barrier coatings are directed toward correlating such models to measured engine performance.

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