US3866417A

US3866417A - Gas turbine engine augmenter liner coolant flow control system

Info

Publication number: US3866417A
Application number: US331078A
Authority: US
Inventors: David A Velegol
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1973-02-09
Filing date: 1973-02-09
Publication date: 1975-02-18
Anticipated expiration: 1992-02-18
Also published as: JPS49112017A; CA995472A; BE810794A; DE2405840A1; FR2217547A1; GB1458054A; IT1006315B

Abstract

A stabilizing and support system for an augmenter cooling liner of a gas turbine engine is shown to include flange means which divide a cooling plenum formed between the cooling liner and an exhaust duct casing into a plurality of individual coolant chambers such that the pressure differential across the cooling liner can be closely controlled. The stabilizing and support system includes a plurality of stabilizers adapted to mount the cooling liner to the exhaust duct casing in such a manner as to define the cooling plenum therebetween. The stabilizers are captured on their outer ends by a stabilizer guide which permits relative thermal expansion to occur between the cooling liner and the exhaust duct casing, and each of the stabilizer guides is connected to a positioning band which is adapted to be mounted to the inside of the exhaust duct casing. In the preferred embodiment, the positioning band is provided with an integrally formed flange which defines a restricted inlet to each of the cooling chambers.

Description

I United States Patent 1 1 1111 3,866,417

velegol 14 1 Feb. 18, 1975 GAS TURBINE ENGINE AUGMENTER [57] ABSTRACT LINER COOLANT FLOW CONTROL SYSTEM A stabilizing and support system for an augmenter [75] lnventor: David A. Velegol, Colliers, W, Va. cooling liner of a gas turbine engine is shown to inelude flan e means which divide a coolin lenum [73] Abblgnee g-enetrahE-lecmtc Company formed bei ween the cooling liner and an exh au st duct incinnati, Ohio casing into a plurality of individual coolant chambers (22] Filed: Feb. 9, 1973 such that the pressure differential across the cooling liner can be closely controlled. The stabilizing and l Appl' 3314,78 support system includes a plurality of stabilizers adapted to mount the cooling liner to the exhaust duct [52] U.S. Cl 60/261, 60/39.32, 60/3966 casing in such a manner as to define the cooling [51 1 Int. Cl F02c 7/20, F03k 3/10 plenum therebetween. The stabilizers are captured on {58] Field of Search 60/261, 39.65, 39.66, 39.69, their outer ends by a stabilizer guide which permits 60/3972 R, 270 R, 39.32, 39.31 relative thermal expansion to occur between the cooling liner and the exhaust duct casing, and each of the [56] References Cited stabilizer guides is connected to a positioning band UNITED STATES PATENTS which is adapted to be mounted to the inside of the 2,510,645 6/1950 McMahan 60/3932 exhaust duct Casing' t Prelerred.embdimem the 2,974,486 3/1961 Edwards 60/261 x Positioning band is Pmvded "tegrally formed 3,031,844 5/1962 TOmOlOniUS 60/3931 x flange which defines a restricted ihlel 10 each of the 3,540,216 11/1970 Quillevere et a1 60/3972 R cooling chambers. 3,570,241 3/1971 Alexander 3,712,062 4/1968 Nash 60/3965 X Primary ExaminerC. J. Husar Assistant Examiner-Robert E. Garrett Attorney, Agent, or Firm-Derek P. Lawrence; Lee H. Sachs 8 Claims, 3 Drawing Figures PATENTED FEB I 8 1975 add-- GAS TURBINE ENGINE AUGMENTER LINER COOLANT FLOW CONTROL SYSTEM BACKGROUND OF THE INVENTION This invention relates generally to augmented gas turbine engines and, more particularly, to means for controlling and regulating augmenter liner coolant pressure.

The invention herein described was made in the course of or under a contract, or a subcontract thereunder, with the United States Department of the Air Force.

Gas turbine engines generally comprise a compressor for compressing air flowing through the engine, a combustion system in which high energy fuel is mixed with the compressed air and ignited to form a high energy gas stream, and a turbine which includes a rotor portion operatively connected to the compressor to drive the same. Many modern-day gas turbine engines are of the turbofan type in which a second or low pressure compressor is mounted forwardly of the high pressure compressor and is driven by a second turbine mounted downstream of the first turbine. The low pressure compressor or fan presents an additional stage of compression and, in addition, is normally of a larger diameter than the high pressure compressor. The turbofan engine is therefore capable of flowing a much larger mass of air, thereby greatly increasing the thrust output of the engine.

An additional known method of increasing the thrust output of the engine is to provide the engine with an augmenter. In such an engine, additional fuel is injected into an exhaust duct formed downstream of the second turbine and is ignited to provide an additional high energy gas stream, which is ejected through an exhaust nozzle to provide high energy thrust output from the engine. In the case of turbofan engines, it is also known to mix the fan airflow with core engine airflow prior to supplying the mixed flow with additional fuel by means of the augmenter.

The augmenter system is normally located within the exhaust duct of the engine and, in most cases, some means must be provided for protecting the exhaust duct from the extremely high temperatures associated with the gas flow within the augmenter. One common means for providing this protection is to position a cooling liner within the exhaust duct and to pass cooling air between the liner and the exhaust duct. Openings or slots are positioned within the cooling liner such that the cooling air may flow through these openings and form a film of coolant on the inside of the cooling liner.

A basic problem confronting the designer of such an augmenter liner consists of the regulation of the coolant flow and the minimization of liner pressure loading. In order to use the minimum amount of cooling air for maximum engine performance, a simple and reliable means is needed to regulate the coolant flow through the openings or slots in the cooling liner. The problem is complicated by the fact that the pressure in the plenum which surrounds the cooling liner is essentially constant along the entire axial length of the liner while the static pressure of the combustion gas inside the liner decreases axially due to the acceleration of the gas as its temperature increases due to operation of the augmenter. This condition not only results in significant pressure differentials across the liner but also in a varying pressure differential from the upstream to the downstream end of such liner.

In order to offset the increasing pressure differential, designers have, in the past, attempted to maintain an essentially constant coolant flow by varying the size of the openings and slots within the liner. In order to offset the higher pressure loading at the aft end of the liner, various means such as reinforcing rings or additional mounting points and hangers have been suggested. Obviously, the additional complexity involved in varying the size of the coolant openings can easily increase the manufacturing costs of such a liner. In addition, the requirements for heavy stabilizing rings and- /or additional mounting hardware may unduly increase the weight of the overall system.

SUMMARY OF THE INVENTION It is an object of the present invention, therefore, to overcome the above-mentioned problems and to provide means for regulating the coolant flow to a plenum surrounding a gas turbine engine augmenter cooling liner such that a relatively constant pressure differential can be achieved along the entire axial length of the liner. It is an additional object of this invention to provide such a coolant regulating means which yields relatively uniform coolant flow through openings provided in the liner without the necessity of varying the size of such openings along the axial length of the liner.

Briefly stated, the above and similarly related objects are attained in the present instance by providing a cooling liner stabilization and mounting system which includes means which divide the coolant plenum surrounding the liner into a number of individual chambers. The flow into each of these chambers is con trolled by means of a flange associated with the mount ing system which acts as a restriction in the flow path and also divides the plenum into the previouslymentioned individual chambers. By varying the size of the flange, the pressure within the individual chambers can be readily controlled and the pressure differential across the liner can thereby readily be regulated.

DESCRIPTION OF THE DRAWING While the specification concludes with a series of claims which particularly point out and distinctly claim the subject matter which Applicant regards as his invention, a clear understanding of the invention will be obtained from the following detailed description, which is given in connection with the accompanying drawing,

in which:

FIG. 1 is a schematic, axial, cross-sectional view of a gas turbine engine incorporting the present inventive regulating means;

FIG. 2 is an enlarged, partial view of the inventive cooling liner stabilization and pressure regulation means of FIG. 1; and

FIG. 3 is a graphical plot showing the pressure differential along the axial length of a cooling liner incorporating the present inventive means.

DESCRIPTION OF A PREFERRED EMBODIMENT Referring to the drawing wherein like numerals correspond to like elements throughout, attention is directed initially to FIG. 1 wherein a gas turbine engine 10 of the mixed flow turbofan type is shown to include a core engine 12 which includes a fan turbine 14 for driving a plurality of fan blades 15 mounted on a shaft 16. The fan blades are located within an inlet 17 formed by an outer or fan casing 18 which surrounds the entire gas turbine engine 10. The fan casing 18 operates with a core engine casing 20 to define

parallel flow paths

22 and 23.

Air entering the flow path 23 is compressed by means of a compressor 24 and is mixed with fuel in combustor 26. Fuel is delivered to the combustor 26 by means of a plurality of fuel injection points 27 from fuel tubes 28 which extend through the flow path 22. The resultant high energy gas stream exits the combustor 26 and drives a turbine 30 which, in turn, drives the compressor 24 by means of a shaft 31.

As further shown in FIG. 1, air flowing through the outer or fan flow path 22 and air exiting the core engine 12 flow through a mixer 32, which operates to mix the two separate flow paths. The mixed flow path is then acted upon by an augmenter 34, which consists of a plurality of fuel injectors 36. The resultant fuel/air mixture in the augmenter 34 is ignited by means of a suitable igniter (not shown), flows through an exhaust duct 40, and thereafter provides an additional propulsive force by exiting through an exhaust nozzle 42.

The exhaust duct is located at the downstream end of the fan casing 18 and is shown in FIG. 1 to include an exhaust duct casing 44 and a cooling air liner which is generally designated by the numeral 46. The cooling liner 46 is spaced radially inwardly from the exhaust duct casing 44 and defines an annular coolant flow path 48 having an inlet 50 formed by a forward lip 52 at the upstream end of the cooling liner 46.

As is well known in the art, the cooling liner 46 includes a plurality of openings or slots 54 adapted to deliver cooling air from the passageway 48 to the inside of the liner 46. The coolant flowing through the open ings 54 provides a film of cool air on the inside of the liner 46 thereby protecting both the liner 46 and the surrounding exhaust duct casing member 44 from the high temperatures associated with the operation of the augmenter 34.

The operation of the engine 10 is well known and will be discussed only briefly. Air flows through the inlet 17 and is acted upon by the fan blades 15. A first portion of this pressurized air flows through the fan flow path 22, while a second portion flows through the core engine flow path 23 and is acted upon by the compressor 24. A high energy gas stream is generated by the combustor 26 and drives the high pressure turbine 30 and low pressure turbine 14, which, in turn, drive the core engine compressor 24 and the fan 15. Air exiting the low pressure turbine 14 and air flowing through the fan flow path 22 are mixed within the mixer 32 and the mixed flow is delivered to the region of the augmenter 34. A resultant fuel/air mixture generated by the augmenter 34 is ignited to provide an additional propulsive force by exiting through the exhaust nozzle 42.

A portion of the air flowing through the fan flow path 22 flows through the inlet 50 and, thus, through the coolant passageway 48. This cooling air thereafter flows through the openings 54 and forms a film on the inside of the cooling liner 46 thereby protecting the liner 46 and the surrounding exhaust duct casing 44 from the high gas temperatures associated with operation of the augmenter 34.

The gas turbine engine 10 described above is typical of many present-day augmented turbofan engines and has been described solely to place the present invention in proper perspective. As will become clear to those skilled in the art, the present invention will be applicable to other types of gas turbine engines and, therefore, the engine 10 is merely meant to be illustrative.

Referring now to FIGS. 2 and 3, the gas turbine engine augmenter cooling liner 46 and its associated pressure regulating system is shown in greater detail. As shown in FIG. 2, the cooling liner 46 is mounted to the exhaust duct 44 by means of a plurality of stabilizer assemblies 56, the details of which are shown and claimed in application Ser. No. 328,769, filed in the name of D. O. Nash et al. and assigned to the same assignee as the present invention now US. Pat. No. 3,826,088. As described in the Nash et al. application, the stabilizer assembly 56 includes a circumferential row of stabilizers 58, each of which are captured within a stabilizer guide 60. The stabilizer guides 60 are, in turn, connected to the inner surface of a positioning band 62, which is adapted to be connected to the inner wall surface of the exhaust duct 44.

The positioning band 62 includes a plurality of holes 64 which are aligned with threaded openings associated with capture nuts 68 mounted on each of the stabilizer guides 60. A plurality of bolts 70 fit through openings 72 formed within the exhaust duct casing 44 and are threadably received within the nuts 68, thereby firmly attaching the positioning band 62 and thus the cooling liner 46 to the exhaust duct casing 44.

As more fully described in the Nash et al application, the stabilizers 58 are adapted to define radial height 71 of the annular coolant passageway 48. As discussed briefly above, tests with previous augmenters have shown that the pressure of the coolant within the passageway 48 is relatively constant along the entire axial length of the liner 46. The static pressure of the combustion gas within the exhaust duct, i.e. inside the liner 46, decreases axially due to the acceleration of the gas as its temperature increases due to operation of the augmenter. With the relatively constant pressure within the passageway 48 and the decreasing pressure within the exhaust duct 40, the pressure differential across the liner 46 increases significantly near the aft end of the liner 46. This condition is shown graphically as a dotted line curve 73 in FIG. 5 wherein pressure differential across the liner is plotted as a function of axial length of the liner.

Referring still to FIGS. 2 and 3, the present invention includes means for regulating the pressure level within the coolant passageway 48 such that the pressure differential across the cooling liner can be maintained relatively constant from the upstream end to the downstream end of the liner. The pressure-regulating means comprise a flange 76 formed at either the upstream or downstream end of the positioning band 62. While the flange 76 may be made a separate member, in the embodiment shown in FIG. 2 the flange 76 is formed integrally with the positioning band 62. As shown best in FIG. 3, the flanges 76 associated with each of the stabilizer assemblies 56 act to divide the coolant passageway 48 into separate

annular chambers

78, 80 and 82.

The flange 76 is sized so as to provide a gap 84 between an inner edge 86 of the flange 76 and the outer wall of the cooling liner 46. The gap 84 acts as a restriction in the flow path for the cooling air and acts to control the pressure of the coolant air within the

chambers

78, 80 and 82. The gap 84 may be made of various sizes depending upon the desired pressure within the

chambers

78, 80 and 82. In certain applications, the gap 84 may be the same for each of the stabilizer assemblies 56, while in other applications the gap size may vary from mounting assembly to mounting assembly. In either case, the gap 84 can be utilized to provide a pressure within the

chambers

78, 80 and 82 such that the average pressure differential across the liner is a minimum in each chamber. This minimum is limited by the least value of pressure differential which will produce the required radially inward cooling flow. This situation is plotted as solid line curve 84 in FIG. 3 with the least desired pressure differential lying below the point A. As shown therein, the pressure differential across the liner 46 may vary slightly from the upstream to the downstream end of each of the

chambers

78, 80 and 82, but the net effect of the invention system is to provide a much lower overall pressure differential from the upstream end of the liner to the downstream end thereof. With such a result, the liner can be made with uniformly sized slots or openings 54 and the need for heavy reinforcing rings at the downstream end thereof becomes non-existent.

It should be obvious to those skilled in the art that slight variations could be made in the structural elements shown and described above without departing from the broader inventive concepts disclosed herein. For example, the flange 76 could be positioned at either end of the positioning band 62. Similarly, the flange 76 could be formed integrally with, or connected to, the stabilizing guides 60 instead of the positioning band 62. Likewise, the flange 76 could be connected directly to either the exhaust duct casing 44 or the cooling air liner 46. In addition, the stabilizing assemblies 56 are capable of use in combination with more standard mounting techniques such as spacers 86 shown in FIG. 3 near the upstream end of the liner 46. The appended claims are intended to cover these and similar modifications in Applicants broader inventive concept.

What is claimed is:

1. In a gas turbine engine of the type including a compressor, a turbine, a combustion system, an augmenter, an exhaust duct surrounding said augmenter, and a cooling liner positioned within said duct so as to form a cooling plenum therebetween, at least a portion of which liner extends downstream from said augmenter and is adapted to protect said exhaust duct from the high temperature gas generated by said augmenter, the

6 improvement comprising:

flange means situated within said cooling plenum and dividing said cooling plenum into at least two individual chambers, said flange means occupying a predetermined radial space in said cooling plenum and providing a restricted inlet between said chambers such that the average pressure differentials across the liner in each of said chambers are substantially equal.

2. The improved gas turbine engine of claim 1 wherein said liner includes a plurality of stabilizer assemblies mounted between said liner and said exhaust duct in such a manner as to define said coolant plenum and each of said stabilizer assemblies includes a plurality of stabilizers, a like number of stabilizer guides, one of which partially surrounds each of said stabilizers, and a positioning band surrounding and mounted to each of said stabilizer guides and to said exhaust duct, and said flange means comprises a flange formed integrally with said positioning band and extending toward said cooling liner.

3. The improved gas turbine engine of claim 1 wherein said inlet between said chambers is sized to provide a relatively constant average pressure differential in each of said chambers along the axial length of said liner.

4. The improved gas turbine engine of claim 2 wherein said cooling liner includes at least two of said stabilizer assemblies such that said plenum is divided into at least three of said chambers.

5. The improved gas turbine engine of claim 4 wherein said cooling liner includes a plurality of relatively equally sized coolant holes spaced along at least a portion of the axial length thereof.

6. The improved gas turbine engine recited in claim 1 wherein said inlets are sized so as to provide a pressure differential across said cooling liner ranging between 0 and 0.5 A P, where A P is defined as pressure in said chamber-pressure in said duct.

7. The improved gas turbine engine recited in claim 6 wherein said pressure differential ranges between 0 and 0.5 A P in each of said chambers.

8. The improved gas turbine engine of claim 4 wherein said restricted inlets are sized and said flange means are axially spaced from one another such that the pressure differential across said cooling liner is reduced.

Claims

1. In a gas turbine engine of the type including a compressor, a turbine, a combustion system, an augmenter, an exhaust duct surrounding said augmenter, and a cooling liner positioned within said duct so as to form a cooling plenum therebetween, at least a portion of which liner extends downstream from said augmenter and is adapted to protect said exhaust duct from the high temperature gas generated by said augmenter, the improvement comprising: flange means situated within said cooling plenum and dividing said cooling plenum into at least two individual chambers, said flange means occupying a predetermined radial space in said cooling plenum and providing a restricted inlet between said chambers such that the average pressure differentials across the liner in each of said chambers are substantially equal.

6. The improved gas turbine engine recited in claim 1 wherein said inlets are sized so as to provide a pressure differential across said cooling liner ranging between 0 and 0.5 Delta P, where Delta P is defined as pressure in said chamber-pressure in said duct.

7. The improved gas turbine engine recited in claim 6 wherein said pressure differential ranges between 0 and 0.5 Delta P in each of said chambers.