Document
[111 4,083,181
United States Patent [191
p ~ ] Apr. 11, 1978
Adamson
3,908,362 9/1975 Szydlowski ......................... 60/39.53 GAS TURBINE ENGINE WITH 3,927,958 2/1975 Quinn .................................. 60/39.52 RECIRCULATING BLEED FOREIGN PATENT DOCUMENTS Arthur P . Adamson, Cincinnati, Ohio Inventor:
6/1956 Canada ........................... 60/39.09 D
527,041 Assignee: The United States of America as 3/1949 Italy .................................... 60/39.52 446,244 represented by the Administrator of 158,752 11/1962 U.S.S.R. .............................. 60/39.52 the National Aeronautics and Space Washington, D.C.
Administration, Primary Examiner-Robert E. Garrett Attorney, Agent, or Firm-Derek P. Lawrence; Robert Appl. NO.: 695,513 C. Lampe, Jr.; John R. Manning Filed: Jun. 14,1976 [571 ABSTRACX
Int. c1.Z ................................................ Fo2c 3/00
A method of reducing carbon monoxide and unburned
U.S. c1. ...................... : .......... 60/39.52; 60/226 R
hydrocarbon emissions i n a gas turbine engine by bleed- Field of Search ............... 60/39.52, 39.53, 226 R, ing hot air from the engine cycle and introducing it 60/39.09 D back into the engine upstream of the bleed location and References Cited upstream of the combustor inlet. As this hot inlet air is U.S. PATENT DOCUMENTS recvcled, the combustor inlet temperature rises rapidlv - . - . r
Clark et al. ......................... 60/39.52 at a engine levdin most
2,404,275 7/1946 this will reduce carbon monoxide and unburned hydro-
Meyer ................................. 60/39.52
2,599,410 6/1952 Torell .................................. 60/39.53 carbon emissions significantly. The preferred locations 2,863,282 12/1958
Haltenkrger ...................... 60139.52
2,906,092 9/1959 for hot air extraction are at the compressor discharge or 3,123,283 3/1964 ................................. 60/39.09 D from the turbine, whereas the preferred Hamla 60/39.52 .................................
3,754,393 8/1973 location is at the compressor inlet.
Amann ................................ 60/39.52
3,785,145 1/1974
Handa ................................. 60/39.52
3,792,581 2/1974 3 C l a i m s , 4 Drawing Figures
8/1974 Ainsworth ......................... 60/226 R
3,830,058
U.S. Patent
April 11, 1978
4,083,181
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location is at least as far upstream as the combustor GAS TURBINE ENGINE WITH RECIRCULATING inlet. Thus, as the hot air is recycled the combustor inlet BLEED temperature rises rapidly for a given engine thrust level so as to reduce CO and HC exhaust emissions.
The invention herein described was made in the per- 5 In the preferred embodiment,a conduit is provided to formance of work under a NASA contract and is sub- transfer hot air from a source such as the compressor ject to the provisions of Section 3 0 5 of the National discharge, combustor discharge or turbine discharge Aeronautics and Space Act of 1958, Public Law 85-568 upstream to the compressor inlet, for example. A valve (72 Stat. 435; 42 USC 2457). within the conduit, and operated by means of a signal 1 0 from the engine fuel control system or power lever, is BACKGROUND OF THE INVENTION provided to control the rate of recirculation of the The present invention pertains to gas turbine engines heated air. Alternatively, it may be possible to dispense and, more particularly, to a method of operating same with such a control valve by finding a combination of to reduce carbon monoxide and unburned hydrocarbon hot air sources and re-entry locations which would 15 permit the bleed air to flow at low power settings and emissions.
not at high power settings. In such event, a simple The present era of environmental awareness has check valve would preclude reverse circulation.
spurred governmental regulations limiting the permissi- ble exhaust emissions from gas turbine engines. Some of BRIEF DESCRIPTION OF THE DRAWINGS the more severe requirements relate to carbon monox- ide (CO) and unburned hydrocarbon (HC) emissions. 20 While the specification concludes with claims partic- These emissions have traditionally been the greatest at ularly pointing out and distinctly claiming the subject matter which is regarded as part of the present inven- ground idle conditions where the combustor inlet tem- tion, it is believed that the invention will be more fully perature and pressure, and the combustor fuel-to-air understood from the following description of the pre- ratio, are relatively low.
As gas turbine powered aircraft are designed for 25 ferred embodiment which is given by way of example operation from shorter runways, the emissions problem with the accompanying drawings in which: will become more acute. The reason is that short-field FIG. 1 represents a schematic, partial cross-sectional view of a gas turbine engine incorporating the subject aircraft must be overpowered (Le., higher installed invention; thrust-to-aircraft weight ratio) compared to the more conventionaltake-off and landing aircraft. For example, 30 FIG. 2 is an enlarged cross-sectional view of a por- tion of the gas turbine engine of FIG. 1; during taxi operation the engine power setting must be FIG. 3 is a gas turbofan engine partial cross-sectional reduced abnormally to avoid overloading the aircraft schematic, similar to FIG. 1, depicting an alternative brakes, particularly on icy runways. As the engine throttle is pulled back to this abnormal position, the embodiment of the present invention; and 35 FIG. 4 is an enlarged cross-sectional view similar to combustor inlet temperature drops (due to lower work FIG. 2 depicting an alternative embodiment of a por- input of the compressor) resulting in inefficient burning and increased exhaust emission levels. A similar condi- tion of the present invention.
tion exists during the landing cycle if the aircraft main- DESCRIPTION OF THE PREFERRED tains a holding pattern, since there again the power EMBODIMENT level must be abnormally low (on a percentage thrust 40 basis) due to the high installed thrust level. Referring to the drawings wherein like numerals correspond to like elements throughout, reference is The problem is further compounded, however, since not only does the low combustor inlet temperature first directed to FIG. 1 wherein an engine depicted generally at 10 and embodying the present invention is result in increased exhaust emissions, but it also de- diagrammatically shown. This engine may be consid- grades the aircraft anti-icing system effectiveness. Some 45 ered as comprising generally a core engine 12, a fan aircraft and engine surfaces are normally heated by air assembly 14 including a stage of fan blades 16, and a fan bled from the combustor inlet and if this air is too cool turbine 18 which is interconnected to the fan assembly the heating process does not function properly.
14 by shaft 20. The core engine 12 includes an axial flow SUMMARY OF THE INVENTION 50 compressor 22 having a rotor 24. Air enters inlet 26 and Accordingly, it is the primary object of the present is initially compressed by fan assembly 14. A first por- invention to provide a method of operating a gas tur- tion of this compressed air enters the fan bypass duct 28 bine engine in order to reduce CO and HC emissions at defined, in part, by core engine 12 and the circumscrib- low power settings. ing fan nacelle 30 and discharges through a fan nozzle is a further object of the present invention to pro- 55 32. A second portion of the compressed air enters inlet It 34, is further compressed by the axial flow compressor vide an improved gas turbine engine having reduced 22 and is then discharged to a combustor 36 where fuel CO and HC emissions at low power settings.
is burned to provide high energy combustion gases These and other objects and advantages will be more which drive a turbine 38. The turbine 38, in turn, drives clearly understood from the following detailed descrip- the rotor 24 through a shaft 40 in the usual manner of a tion, drawings and specific examples, all of which are 60 gas turbine engine. The hot gasses of combustion then intended to be typical of rather than in any way limitng pass to and drive the fan turbine 18, which, in turn, to the scope of the present invention.
Briefly stated, the above objects are accomplished in drives the fan assembly 14. A propulsive force is thus a gas turbine engine wherein hot air is bled from the obtained by the action of the fan assembly 14 discharg- ing air from the fan bypass duct 28 through the fan engine at a first location and reintroduced back into the 65 nozzle 32 and by the discharge of combustion gases engine at a second location, with the necessary con- from a core engine nozzle 42 defined, in part, by plug straints that the temperature of the air at the first loca- 44. The foregoing description is typical of many pres- tion exceedsthat of the second and wherein the re-entry
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ent-day gas turbine engines and is not meant to be limit- required, and the means necessary for extracting the ing, as it will become readily apparent from the follow- bleed air and reintroducing it into the compressor inlet.
ing description that the present invention is capable of The obvious choice for the bleed extraction location application to any gas turbine engine and is not neces- in contemplation of modifying existing engines is at the sarily restricted to gas turbine engines of the turbofan 5 compression discharge location (45 of FIG. 1 ) as dis- variety. The foregoing description of the operation of cussed hereinabove, by utilizing extraction ports al- the engine as depicted in FIG. l is, therefore, merely ready in the engine for customer purposes such as air- meant to be illustrative of one type of application for the craft cabin pressurization or anti-icing. However, sev- present invention. eral other sources may be tapped for the hot air depend- For most gas turbine engine combustors it has been 10 ing upon the amount of flow and temperature rise de- found that the amount of gas emissionsat idle (or below sired. Means for bleeding this hot air are indicated by idle) engine operating conditions can be reduced by the dotted lines feeding conduit 48 in FIG. 1 . Specifi- increasing the temperature of the air entering the com- cally, they include compressor interstagebleed 50, com- bustor. In the present invention, the temperature of the bustor inlet bleed 52, combustor discharge bleed 54, air entering the combustor is increased by recycling the 15 turbine interstage bleed 56 and turbine discharge bleed heated air in any of several ways now to be described. 58. Clearly, extraction from the turbine area where the Generally, a portion of the air is bled from the engine at air is at a much higher temperature will provide a much a first location and reintroduced back into the engine at, greater increase in compressor inlet and exhaust tem- a second location subject to two constraints: perature for a given bleed flow rate. For example, if (1) the temperature of the air (the motive fluid) pass- 20 combustor discharge bleed air at a temperature of 800" ing through the engine must be higher at the bleed F (426.7" C) were used in the previous example, the source (the first location) than at the re-entry location amount of bleed air required to increase the compressor (the second location); and inlet temperature by 50" F (27.8" C) would be only (2) the re-entry location must be at least as far up- approximately 4.5 percent of the total air available pass- stream as the inlet to the combustor. 25 ing through the engine. Note also that the cycle re- For example, consider the arrangement of FIG. 1 matching effect will be different for each extraction wherein air is bled from the discharge 45 of compressor location.
22 and routed by means of conduits 46, 48 to the core FIG. 1 depicts the obvious choice for the re-entry compressor inlet 34 where it is reintroduced back into location, at the core compressor inlet 34. While the the primary flow stream. Since the temperature at the 30 concept is depicted only schematically in FIG. 1 , the compressor discharge is greater than that at the inlet by geometry of the reintroducing means may vary depend- virtue of the work addition through the compressor, the ing upon individual engine differences and design pref- average compressor inlet temperature is increased. erences. For example, FIG. 2 shows one possible ar- When the engine cycle is rebalanced in the known man- rangement wherein the bleed air is ducted into the flow ner to supply a specific idle or subidle thrust, the net 35 splitter 60 separating the core inlet duct 34 from the fan result is an increase in combustor inlet temperature of an bypass duct 28. Therein, the bleed air is fed into a ple- amount in excess of the increase in inlet temperature. num 62 within the splitter and ejected therefrom This, in turn, reduces CO and HC emissions signifi- through means such as representative apertures 64 to mix with the incoming air of inlet 34. Alternatively, as cantly.
A simple estimate of the effectiveness of the concept 40 is best shown in FIG. 4, the hot bleed air could be on a commercially available high bypass-ratio gas tur- routed from the bleed location to the interior 65 of one bofan engine indicates that a 50" F (27.8" C) increase in of a plurality of hollow frame struts 66 (which typically core compressor inlet temperatu?e at a 5 percent thrust support the splitter 60 in its proper spacial relationship idle condition will increase the combustor inlet temper- with the core engine 12) by means of a conduit 67 and ature by 75" F (41.7" C). This, in turn, decreases the CO 45 ejected therefrom through a plurality of apertures 69. In emissions by approximately 28 percent, even after ac- any eyent, the design should provide for the suitable counting for an increased fuel flow of about 4 percent mixing of the gases with the primary core engine stream required to rebalance the cycle. The magnitude of the and such mixing should occur early in the compression compressor discharge bleed required to raise the com- process.
pressor inlet temperature by 50" F is about 14 percent of 50 Means such as valve 68 is provided in conduit 48 to the total air passing through the core compressor. It is permit the hot bleed air to be reintroduced into the to be noted that these estimates do not include the effect compressor inlet airstream only at the abnormally low of bleed on the cycle, but just include the inlet tempera- power settings discussed hereinabove. Typically, it is ture effect. The effect of compressor bleed by itself anticipated that such a valve would be controlled without recirculation is to raise turbine inlet tempera- 55 through the main engine fuel control means 70 which, ture and the combustor fuel-to-air ratio required to in turn, is controlled by the pilot through throttle quad- supply a given level of idle thrust. This will further rant 7 2 . The particular type of valve and its method of contribute to reduced emissions and the two effects will control are well within the capability of engine design- complement each other. Thus, the improvement in idle ers and the details need not be elaborated herein. One emissions achieved by raising the compressor inlet tem- 60 example which may be adapted to the present invention perature is in addition to the effect of compressor dis- is the torque and power sensing and control system charge bleed. taught in US. Pat. No. 3,106,062 which is assigned to In principle, the magnitude of the increases in com- the assignee of the present invention and the subject pressor inlet and discharge temperatures can be made matter of which is incorporated herein by reference.
any reasonable value depending on the particular en- 65 In some cases it may be possible to dispense with gine involved, the power setting in consideration, and valve 68 by finding a hot a i r source which is at a higher practical considerations such as maximum temperature pressure than the re-entry point at low power settings, limits of the compressor inlet, the size of the ducting and at a low pressure at high power settings. Such an
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arrangement is depicted in FIG. 3 where bleed a i r is nected to said compressor, the improvement compris- extracted downstream of the turbine at 73 and reintro- ing: d u d in the early compressor stages at 7 4 . Such an means for bleeding a portion of the motive fluid from arrangement is possible since the pressure level in the a first engine location at a relatively high tempera- early compressorstages is subatmosphericat low power 5 ture; settings,and any bleed flow would naturally occur from means for reintroducing the Portion at substantially right to left at FIG. 3 , whereas the flow would reverse the same high temperature back into the engine itself and flow rearward (left to right) at higher power motive fluid stream at a second location upstream settings. This undesirable rearward flow may be pre- of the combustor wherein it is mixed with motive fluid at a relatively lower temperature with respect vented by such as a check valve 76, if 10 to the bleed portion; desired.
Therefore, a method has been provided for reducing valve means in serial flow relationship between the bleed meeans and the reintroducing means; and CO and HC emissions in a gas turbine engine by bleed- control m e w responsive to engine power setting for ing a portion of the motive fluid (for example, &) from positioning said valve means to limit the bleed flow 15 a first location and reintroducing it back into the engine to low engine power operation.
motive stream at a second location as long as the bleed 2, In a gas turbine engine having a compressor for air is at a higher temperature than the motive stream at pressurizing a motive fluid stream, a combustor down-
the reentry location, and = long as the re-entry l°Ca-
stream of said compressor and a turbine drivingly con-
' at least = far as the inlet*
2 4 ) nected to said compressor, the improvement compris- It will become obvious to one skilled in the art that ing: certain changes and can be made to the means for bleeding a portion of the motive fluid from above-described invention without departing from the a first engine location at a relatively high tempera- broad inventive concepts thereof. For example, while ture; the Of flow has been Only sche- 25 means for reintroducing the portion at substantially matically in FIGS. 1 - 3, it will be recognized that such the same high temperature back into the engine piping and ducting may be either internal or external to motive fluid stream at a second location upstream the engine while still being within the scope of the of the combustor wherein it is mixed with motive present invention. Furthermore, the present invention is fluid at a relatively lower temperature with respect applicable to other types of gas turbine engines includ- 3 0 to the bleed portion; and ing, but not limited to, those of the turbojet and boosted check valve means in serial flow relationshipbetween turbofan varieties. It is intended that the appended the bleed means and the reintroducing means; and claims cover these and all other variationsin the present wherein the fvst engine location is at a higher pres- invention's broader inventive concepts.
sure than the second location at low engine power Having thus described the invention, what i s claimed 35 operation, and at a lower pressure than the second as novel and desired to be secured by Letters Patent of location at higher engine power operation.
the United States is: 3. The engine as recited in claim 2 wherein said com- 1. In a gas turbine engine having a compressor for pressor is a multistage compressor and said second loca- pressurizing a motive fluid stream, a combustor down- tion is in the early compressor stages.
* * * * *
stream of said compressor and a turbine drivingly con- 40