CA1185101A - Drag-reducing nacelle - Google Patents
Drag-reducing nacelleInfo
- Publication number
- CA1185101A CA1185101A CA000367228A CA367228A CA1185101A CA 1185101 A CA1185101 A CA 1185101A CA 000367228 A CA000367228 A CA 000367228A CA 367228 A CA367228 A CA 367228A CA 1185101 A CA1185101 A CA 1185101A
- Authority
- CA
- Canada
- Prior art keywords
- aft
- fan
- cowl
- nozzle
- engine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
Links
- 230000003247 decreasing effect Effects 0.000 claims abstract description 7
- 239000007789 gas Substances 0.000 claims description 16
- 238000011144 upstream manufacturing Methods 0.000 claims description 10
- 239000000543 intermediate Substances 0.000 claims 8
- 238000007599 discharging Methods 0.000 claims 1
- 238000010276 construction Methods 0.000 abstract description 6
- 230000001627 detrimental effect Effects 0.000 abstract 1
- 239000003570 air Substances 0.000 description 19
- 230000003068 static effect Effects 0.000 description 14
- 230000007423 decrease Effects 0.000 description 7
- 230000000694 effects Effects 0.000 description 6
- 239000012080 ambient air Substances 0.000 description 3
- 230000009467 reduction Effects 0.000 description 3
- 230000000875 corresponding effect Effects 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 230000008569 process Effects 0.000 description 2
- 230000002829 reductive effect Effects 0.000 description 2
- 230000002411 adverse Effects 0.000 description 1
- 238000013459 approach Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 230000005465 channeling Effects 0.000 description 1
- 230000003993 interaction Effects 0.000 description 1
- 230000000670 limiting effect Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C23/00—Influencing air flow over aircraft surfaces, not otherwise provided for
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D29/00—Power-plant nacelles, fairings, or cowlings
- B64D29/02—Power-plant nacelles, fairings, or cowlings associated with wings
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
- B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/52—Nozzles specially constructed for positioning adjacent to another nozzle or to a fixed member, e.g. fairing
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C2230/00—Boundary layer controls
- B64C2230/20—Boundary layer controls by passively inducing fluid flow, e.g. by means of a pressure difference between both ends of a slot or duct
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/10—Drag reduction
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Aviation & Aerospace Engineering (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Exhaust Silencers (AREA)
- Control Of Turbines (AREA)
- Jet Pumps And Other Pumps (AREA)
Abstract
DRAG-REDUCING NACELLE
ABSTRACT OF THE DISCLOSURE
A Wing-mounted gas turbofan engine is provided with an exhaust system that reduces airflow drag during subsonic aircraft flight peration. The system reduces drag by turning the engine bypass exhaust stream away from the underside of the airplane wing to lessen its tenddency to reduce air pressure under the wing. Low pressure below the airplane wing has a detrimental effect on wing lift which, in turn, causes greater airflow drag. The exhaust system construction lessens its pressure lowering influence by first inwardly curving the exhaust system exit to turn the bypass flow away from the airplane wing; second, by determining the location of a nozzle throat within the bypass stream to control exit exhaust pressure so as to match outside air pressure and third, by decreasing the diameter of a portion of the engine nacelle just downstream of the bypass exhaust steam exit to provide a flow region for the exhaust that is further from the airplane wing.
ABSTRACT OF THE DISCLOSURE
A Wing-mounted gas turbofan engine is provided with an exhaust system that reduces airflow drag during subsonic aircraft flight peration. The system reduces drag by turning the engine bypass exhaust stream away from the underside of the airplane wing to lessen its tenddency to reduce air pressure under the wing. Low pressure below the airplane wing has a detrimental effect on wing lift which, in turn, causes greater airflow drag. The exhaust system construction lessens its pressure lowering influence by first inwardly curving the exhaust system exit to turn the bypass flow away from the airplane wing; second, by determining the location of a nozzle throat within the bypass stream to control exit exhaust pressure so as to match outside air pressure and third, by decreasing the diameter of a portion of the engine nacelle just downstream of the bypass exhaust steam exit to provide a flow region for the exhaust that is further from the airplane wing.
Description
~L85~
- l - 13DV-7649 DRAG-RE UCING NACELLE
Backgro_nd of the Invention Field of the Invention This invention relates to nacelle and bypass duct construction for wing-mounted gas turbofan aircraft engines.
Summary of the Prior Art . . _ _ . ~
It is well known that lifting forces are produced by an aircraft wing during flight as a resul-t of a pressure difference acting over the wing platform. As the wing passes through a volume of air, relatively high air pressure is developed below the wing and relatively low air pressure is developed above the wing. In general, the greater the difference in pressure between the upper surface and lower surface of the wing, the greater the lift produced by the wing.
It is also well known that as the airplane is more steeply angled, the angle of attack of the wing is increased, and the pressure differences and lift are correspondingly increased. Unfortunately, an increase in angle of attack also has a corresponding effect on aerodynamic drag produced by the wing. Because the angle of attack of the wing is increased to produce more lift, the wing projects a greater frontal area causing the increase in drag.
When an aircraft is tra~elling a-t subsonic speeds, an engine that is positioned beneath the aircraft ~s~
13DV-76~9 wing causes the local wing underside pressures to be lower than what they would be under the same wing without the engine. This localized lowering of the underside pressure results in a decreased pressure differential and reduces the wing lift for a given angle of attac~. Since a given aircraft requires a fixed amount of lift to maintain altitude at a gi-~en cruise velocity, the wing angle of attack must be increased to regain that amount of lift which is lost due to the presence of the engine nacelle. As expected, this increase in -the angle of attack re~uired to offset the lift loss caused by the engine results in another increase in aerodynamic drag. Those skilled in the art commonly refer to this drag produced by the presence of the nacelle under the wing as "interference drag."
Analysis of interference drag has revealed that different engine nacelle shapes may have similar or identical isolated drag by themselves in an airstream, but can have very diferent effects on a wing pressure distribution, and thus create widely differing amounts of interference drag. Further analysis has been directed at understanding these differences and the causes of this interference drag. The results of this analysis indicate that efforts should be directed towards minimizing the effect of engine fan exhaust systems on wing pressure distribution for the purpose of reducing interference drag.
Summary of the Inven-tion It is, therefore, an object of the present invention to reduce interference drag caused by winy-mounted aircraft engines.
It is another object of the present invention to minimi~e the ef~ect of wing-mounted engine fan exhaust systems on the pressure distribution underneath an airplane wing.
It is still another object of the present invention to modify engine nacelle and fan exhaust construction for the purpose of improving the effect of the engine fan exhaust system on air pressure beneath the airplane wing during subsonic aircraft operation.
These and other ojbects will be more fully understood from the drawings and from the following description, all of w~ich are intended to be representative of, rather than in any way limiting on, the scope of the invention.
Briefly, in one embodiment of the present invention, the engine nacelle and bypass duct construction is modified for the purpose of redirecting bypass exhaust air to minimize its influence on wing underside pressures. First, the inner profile of the bypass duct is curved radially inward at its aft end for the purpose of physically turning the bypass flow radially inward and away from the underside of the wing. Second, a nozzle throat is formed within the bypass duct at a more upstream position in respect to previous practice. The throat is located upstream at a particular location to obtain a bypass duct exit pressure that closely matches outside air pressure, so the exhaust doesn't expand and flow in the direction of the wing. Third, the outer diameter of a portion of the nacelle that is located immediately downstream of the bypass duct exit is reduced in diame-ter and curved radially inward to provide a flow region for the exhaust stream at a location that is more distant from the airplane wing.
_scription of the_Drawings While the specification concludes with t~e claims distinctly claiming and particularly pointing out the invention described herein, it is believed that the invention will be more clearly understood by reference to the discussion below in conjunction with the following drawings:
Figure 1 is an elevation view of a prior art ~ing-mounted turbofan engine and its associated exhaust flow stream pattern;
Figure 2 is a graphical representation of local static air pressure (P ) as a function of cross-sectional area (A) in a nozzle or channeled flowpath;
Figure 3 is a view of a prior art gas turbofan engine, partly in cross section and partly broken away, and the engine's fan air bypass ~low stream pattern;
Figure 4 is a view of a gas turbofan engine, partly in cross section and partly broken away, that incorporates the present invention, and the engine's fan air bypass flow stream pattern;
~igure 5 is a cross-sectional view of the gas turbofan engine shown in Figure 3 overlayed wi~h a dashed outline o~ the gas turbofan engine of Figure 4 that incorporates the present lnvention.
Description of the Preferred Embodiment .
Referring now to Figure 1, a conventional wing-mounted gas turbofan engine 10 is shown suspended by a pylon 12 from an airplane wing 1~. An aircraft with the engine and wing arrangement shown in Figure 1 is designed for subsonic operation. The engine 10 is a typical high-bypass turbofan aircraft engine that has an outer covering or nacelle 15 comprising a relatively large radius fan cowl 16 at its upstream or forward portion and a relatively smaller radius core cowl 18 at its downs-tream or aft region. The fan cowl 16 covers a fan section of the engine where rotating fan blades accelerate a large volume of air ~35~
in aft direction. Some of this air that is accelerated by the fan bypasses a turbine section of the engine and is exhausted from an aft section of the fan cowl 16in the region radially surrounding the core cow]. 18.
The remaining portion of the fan air is drawn at inlet 17 into the turbine section of the englne where it is used in the combustive processes to produce engine power. After flowing through the turbine, gases resulting from the combustive processes are exhausted further downstream out the aft end 19 of the core cowl 18.
Analysis has shown that there are at least three major factors that influence the mutual interaction between external subsonic airflow adjacent to the lower surface of the wing 1~ and supersonlc airflow that is discharged from the aft end of the ~an cowl 16. Referri.ng again to Figure 1, a first factor is minimum physical distance, generally indicated by arrow 20, between the wing lower surface and what is referred to as a dividing streamline 22. This dividing streamline is a boundary between fan airflow exhausted from the fan cowl 16, and surrounding ambient airflow that passes around the outside of the fan cowl 16.
This dividing streamline is also known to those skilled in the art as a "slip line" and is shown in its normal position during flight cruise conditions by the wavy line 22.
A second factor is overall pressure ratio of the fan airflow exiting the fan cowl in respec-t to ambient air pressure (PT/F~/Po)- PT/FAN P
the stagnation pressure of the exhaused fan airflow, and P represents the ambient static air pressure.
A third factor is the Mac~ number of the ambient airflow which passes externally around -the fan cowl 16.
The flow of ambient air between the lower surface of the wing 14 and the dividing streamline 22 is similar in some respects to flow of air through a duct of varying cross-sectional area. This changing cross-sectional area creates a "channeling" e-ffect on ambient air flowing between the engine and wing which is similar to the effect caused by a nozzle.
Referring now to Figure 2, variation of local static pressure (PS/Pt) in a channel or nozzle is shown as a function of a cross-sectional flow area that approximates the cross-sectional area between the wing lower surface and the dividing s-treamline 22 in Figure 1. In explanation of Figure 2, A is the local cross-sectional ~rea, A* is a reference throat or minimum area of that "channel" between thc wing and engine, Ps is local static pressure, and Pt is stagnation pressure for a gi-~en flow~ Both 8* and Pt are constan-t for a given flow rate through the duct. The graph shows that when the flow upstream of the throat (A*) is subsonic (M 1.0), a decrease in the duct area causes a decrease in the local static pressure (Ps), and when the upstream flow is supersonic (M 1.0) t an increase in duct area causes a further decrease in static pressure. This behavior is typical of airflow through a nozzle and is well-known among aeronautical and mechanical engineers. The important feature of this physical phenomenon is tha-t a channeled area or no~zle creates a rapid decrease in local static pressure (Ps) as airflow goes from subsonic (M 1.0) to supersonic (M 1.0). This is what occurs between an aircraft wing and an aircraft engine. When static pressure drops because of this nozzle effect in the region below an airplane wing, an adverse e~ect is created on wing lift.
Referring again to Figure 1, the flow between 13D~-7649 the lower surface of -the wing 14 and the dividing streamline 22 behaves ln a manner very similar to the flow through a duct of varying area as described above. Starting at a leading edge of the wing 1~, it can be readily appreciated that the dis-tance between the wing lower surface and the dividing streamline 22 decreases do~n to a minimum value at some axial loca-tion aft of the wing leading edge, generally shown at arro~ 20 in Figure 1. The presence of the engine nacelle 15 and the trailing dividing streamline near the underside of the wing 14 creates this "channel" or "nozzle" wi-th a throat at the location of arrow 20. The magnitude of the pressure reduction, and the magnitude of the loss of lift to the aircraft, is a function of nacelle position and the position lS of the fan jet dividing streamline 22 relative to the wing 1~. The more the streamline 22 "billows out" and approaches the lower wing surface, the greater the reduction in area between the wing 14 and dividing streamline 22, and thus the lower the air pressure under the wing 1~. If the position of the engine nacelle is fixed, the position of the fan jet dividing streamline 22 must be altered to decrease lift loss, allowing the aircraft -to maintain a lower angle of attack and reduce the correspond~ilg induced aerodynamic drag.
There are at least three factors that can be altered by engine designers that have an effect on the shape of the fan jet dividing streamline 22. These are the pressure of fan air exhaust, the shape of the trailing edge of the fan cowl 16, and the s.hape of the outer surface of the core cowl 18.
P~eferring no~ to Figure 3, a portion of the trailing edge of the fan cowl 16 and a portion of the core cowl 18 is shown for the purpose of explaining the influence of these three factors on the fan jet 13DV-~649 dividing streamline 22~ The space between -the aft portion of the fan cowl 16 and a forward portion of the core cowl 18 is called a bypass duct 24. The bypass duct encloses the path taken by the fan air that bypasses -the turbine section of the engine. The lines projected from the bypass duct at the aft tip of the fan cowl 16 are provided for the purpose of showing the influences of initial discharge angle, shown as 26 in Figure 3, and the exit static pressure ratio, on the shape of the dividing s-treamline 22. It can be readily appreciated from the drawing that the larger the initial discharge angle 26, the larger the maximum diameter of the dividing streamline. Similarly, the higher the exit static pressure ratio PE/Po (static pressure at exit/static pressure outside the fan cowl~, the larger the maximum diameter of the dividing streamline. The exit pressure (PE) will affect the dividing streamline because gas exiting at a higher pressure will have a greater tendency to expand radially outwardly into the surrounding airflow.
Finally, the larger the radius of the core cowl 18 relative to the engine centerline, the more the core cowl will physically force the bypass flow radially outward thereby increasing the maximum diameter of the dividing streamline. Because an increase in the maximum diameter of the dividing streamline 22 reduces the flow area between the wing lower surface and the streamline 22, pressure below the wing surface is reduced and there is an induced drag penalty as explained earlier. Any change in the construction of the fan cowl 16, bypass duct 24, and core cowl 18, which would decrease the maximum di~meter of the dividing streamline 22, will have a corresponding beneficial ef~ect on wing lift, thereby decreasing induced drag. This is the object of -the present invention.
Referring now to Figure 4, a cross-sectional view of a turbofan engine 10 is shown that incorporates one embodiment of the present invention. The invention utilizes three separate constructional features that improve the enginels bypass air exhaus-t system to reduce the maximum radius of the dividing streamline 22 and thus reduce drag. First, the trailing edge oE the inner sur~ace of the fan cowl 16 which forms the outer surface of the aft end of the bypass duct 2~, is reformed so that the downstream portion 28 of the fan cowl is curved radially inward for the purpose of directing the fan exhaust flow radially inward in respect to engine centerline. In the embodiment shown in Figure 4, the downstream portion 28 is curved radially inwardly from a position opposite the maximum radius of the core cowl 18 to the end of the bypass duct 24.
The second feature of th~ present invention is a reconstruction of the flow area distribution at the aft end of the bypass duct 24. This is accomplished by movin~ the minimum cross-sectional area or nozzle ~Jthroat ~ upstream or forward of the bypass duct exit so that the throat of the nozzle is not located where the bypass flow exhausts into the surrounding ambien-t air. By moving the nozzle throat forward, -the flow area distribution at the downstream end of -the bypass duct is increasing, thereby forming a converging-diverging nozzle. Because the bypass flow at the throat of the nozzle is choked, the bypass flow in the diverging section of the nozzle expands and looses pressure in the downstream directionO The length of the divergingsection is carefully predetermined such that the pressure at the nozzle exit is approximately equal to the ambient airstream pressure at the exit 13DV-76~9 of the fan cowl 16 during aircraft cruise operation.
This provides an exit static pressure ratio (PE/P~) of approximately 1Ø An exit static pressure ration of 1.0 causes the nozzle discharge angle of ~he bypass flow to be essentially equal to the angle of the inner wall of the fan cowl at 23. If this pressure ratio were greater than 1.0, the discharge angle would be greater than the wall angle, thus causing the jet plume to billow out relative to the b~pass wall angle.
The third feature of the present invention that reduces the maximum diameter of the dividing streamline is a reconstruction of the shape of the conic core cowl 18. ~ssentially, the conic core co~l 18 is provided with a steadily decreasing outer radius from the nozzle throat to the aft end of the core cowl. For a given amount of bypass flow passing over any core cowl at a given pressure ratio, the cowl with a lesser maximum outer radius will generally produce a lesser dividing streamline maximum diameter~
A reduction in core cowl radius provides a flow area for the fan exhaust stream that is closer to the engine 15 centerline and further from the airplane wing. This relocated flow area contributes to the effec-t of relocating the dividing streamline 22 further from the airplane wing 14.
Referring now to Figure 5, the engine and nacelle employing the present invention from Figure ~
is superimposed in dashed outline 29 on the prior art engine and nacelle of Figure 3. The differences in construction of the fan cowl 16, core cowl i8, and the downstream portion of the fan cowl 28, can be readily appreciated. A region that separates the dividing streamlines of the two engines is additionally shown with a cross-hatched section 30. The outer perimeter 32 of this cross-hatched section is the dividing streamline location for an englne employing the present invention.
The difference in proximity to the airplane wing is readily apparent.
While specific embodiments have been described, it will be apparent to those skilled in -the art that various modifications thereto can be made without departing from the scope of the invention~ as recited in the appended claims. The scope of the invention, therefore, is to be derived from the following claims.
- l - 13DV-7649 DRAG-RE UCING NACELLE
Backgro_nd of the Invention Field of the Invention This invention relates to nacelle and bypass duct construction for wing-mounted gas turbofan aircraft engines.
Summary of the Prior Art . . _ _ . ~
It is well known that lifting forces are produced by an aircraft wing during flight as a resul-t of a pressure difference acting over the wing platform. As the wing passes through a volume of air, relatively high air pressure is developed below the wing and relatively low air pressure is developed above the wing. In general, the greater the difference in pressure between the upper surface and lower surface of the wing, the greater the lift produced by the wing.
It is also well known that as the airplane is more steeply angled, the angle of attack of the wing is increased, and the pressure differences and lift are correspondingly increased. Unfortunately, an increase in angle of attack also has a corresponding effect on aerodynamic drag produced by the wing. Because the angle of attack of the wing is increased to produce more lift, the wing projects a greater frontal area causing the increase in drag.
When an aircraft is tra~elling a-t subsonic speeds, an engine that is positioned beneath the aircraft ~s~
13DV-76~9 wing causes the local wing underside pressures to be lower than what they would be under the same wing without the engine. This localized lowering of the underside pressure results in a decreased pressure differential and reduces the wing lift for a given angle of attac~. Since a given aircraft requires a fixed amount of lift to maintain altitude at a gi-~en cruise velocity, the wing angle of attack must be increased to regain that amount of lift which is lost due to the presence of the engine nacelle. As expected, this increase in -the angle of attack re~uired to offset the lift loss caused by the engine results in another increase in aerodynamic drag. Those skilled in the art commonly refer to this drag produced by the presence of the nacelle under the wing as "interference drag."
Analysis of interference drag has revealed that different engine nacelle shapes may have similar or identical isolated drag by themselves in an airstream, but can have very diferent effects on a wing pressure distribution, and thus create widely differing amounts of interference drag. Further analysis has been directed at understanding these differences and the causes of this interference drag. The results of this analysis indicate that efforts should be directed towards minimizing the effect of engine fan exhaust systems on wing pressure distribution for the purpose of reducing interference drag.
Summary of the Inven-tion It is, therefore, an object of the present invention to reduce interference drag caused by winy-mounted aircraft engines.
It is another object of the present invention to minimi~e the ef~ect of wing-mounted engine fan exhaust systems on the pressure distribution underneath an airplane wing.
It is still another object of the present invention to modify engine nacelle and fan exhaust construction for the purpose of improving the effect of the engine fan exhaust system on air pressure beneath the airplane wing during subsonic aircraft operation.
These and other ojbects will be more fully understood from the drawings and from the following description, all of w~ich are intended to be representative of, rather than in any way limiting on, the scope of the invention.
Briefly, in one embodiment of the present invention, the engine nacelle and bypass duct construction is modified for the purpose of redirecting bypass exhaust air to minimize its influence on wing underside pressures. First, the inner profile of the bypass duct is curved radially inward at its aft end for the purpose of physically turning the bypass flow radially inward and away from the underside of the wing. Second, a nozzle throat is formed within the bypass duct at a more upstream position in respect to previous practice. The throat is located upstream at a particular location to obtain a bypass duct exit pressure that closely matches outside air pressure, so the exhaust doesn't expand and flow in the direction of the wing. Third, the outer diameter of a portion of the nacelle that is located immediately downstream of the bypass duct exit is reduced in diame-ter and curved radially inward to provide a flow region for the exhaust stream at a location that is more distant from the airplane wing.
_scription of the_Drawings While the specification concludes with t~e claims distinctly claiming and particularly pointing out the invention described herein, it is believed that the invention will be more clearly understood by reference to the discussion below in conjunction with the following drawings:
Figure 1 is an elevation view of a prior art ~ing-mounted turbofan engine and its associated exhaust flow stream pattern;
Figure 2 is a graphical representation of local static air pressure (P ) as a function of cross-sectional area (A) in a nozzle or channeled flowpath;
Figure 3 is a view of a prior art gas turbofan engine, partly in cross section and partly broken away, and the engine's fan air bypass ~low stream pattern;
Figure 4 is a view of a gas turbofan engine, partly in cross section and partly broken away, that incorporates the present invention, and the engine's fan air bypass flow stream pattern;
~igure 5 is a cross-sectional view of the gas turbofan engine shown in Figure 3 overlayed wi~h a dashed outline o~ the gas turbofan engine of Figure 4 that incorporates the present lnvention.
Description of the Preferred Embodiment .
Referring now to Figure 1, a conventional wing-mounted gas turbofan engine 10 is shown suspended by a pylon 12 from an airplane wing 1~. An aircraft with the engine and wing arrangement shown in Figure 1 is designed for subsonic operation. The engine 10 is a typical high-bypass turbofan aircraft engine that has an outer covering or nacelle 15 comprising a relatively large radius fan cowl 16 at its upstream or forward portion and a relatively smaller radius core cowl 18 at its downs-tream or aft region. The fan cowl 16 covers a fan section of the engine where rotating fan blades accelerate a large volume of air ~35~
in aft direction. Some of this air that is accelerated by the fan bypasses a turbine section of the engine and is exhausted from an aft section of the fan cowl 16in the region radially surrounding the core cow]. 18.
The remaining portion of the fan air is drawn at inlet 17 into the turbine section of the englne where it is used in the combustive processes to produce engine power. After flowing through the turbine, gases resulting from the combustive processes are exhausted further downstream out the aft end 19 of the core cowl 18.
Analysis has shown that there are at least three major factors that influence the mutual interaction between external subsonic airflow adjacent to the lower surface of the wing 1~ and supersonlc airflow that is discharged from the aft end of the ~an cowl 16. Referri.ng again to Figure 1, a first factor is minimum physical distance, generally indicated by arrow 20, between the wing lower surface and what is referred to as a dividing streamline 22. This dividing streamline is a boundary between fan airflow exhausted from the fan cowl 16, and surrounding ambient airflow that passes around the outside of the fan cowl 16.
This dividing streamline is also known to those skilled in the art as a "slip line" and is shown in its normal position during flight cruise conditions by the wavy line 22.
A second factor is overall pressure ratio of the fan airflow exiting the fan cowl in respec-t to ambient air pressure (PT/F~/Po)- PT/FAN P
the stagnation pressure of the exhaused fan airflow, and P represents the ambient static air pressure.
A third factor is the Mac~ number of the ambient airflow which passes externally around -the fan cowl 16.
The flow of ambient air between the lower surface of the wing 14 and the dividing streamline 22 is similar in some respects to flow of air through a duct of varying cross-sectional area. This changing cross-sectional area creates a "channeling" e-ffect on ambient air flowing between the engine and wing which is similar to the effect caused by a nozzle.
Referring now to Figure 2, variation of local static pressure (PS/Pt) in a channel or nozzle is shown as a function of a cross-sectional flow area that approximates the cross-sectional area between the wing lower surface and the dividing s-treamline 22 in Figure 1. In explanation of Figure 2, A is the local cross-sectional ~rea, A* is a reference throat or minimum area of that "channel" between thc wing and engine, Ps is local static pressure, and Pt is stagnation pressure for a gi-~en flow~ Both 8* and Pt are constan-t for a given flow rate through the duct. The graph shows that when the flow upstream of the throat (A*) is subsonic (M 1.0), a decrease in the duct area causes a decrease in the local static pressure (Ps), and when the upstream flow is supersonic (M 1.0) t an increase in duct area causes a further decrease in static pressure. This behavior is typical of airflow through a nozzle and is well-known among aeronautical and mechanical engineers. The important feature of this physical phenomenon is tha-t a channeled area or no~zle creates a rapid decrease in local static pressure (Ps) as airflow goes from subsonic (M 1.0) to supersonic (M 1.0). This is what occurs between an aircraft wing and an aircraft engine. When static pressure drops because of this nozzle effect in the region below an airplane wing, an adverse e~ect is created on wing lift.
Referring again to Figure 1, the flow between 13D~-7649 the lower surface of -the wing 14 and the dividing streamline 22 behaves ln a manner very similar to the flow through a duct of varying area as described above. Starting at a leading edge of the wing 1~, it can be readily appreciated that the dis-tance between the wing lower surface and the dividing streamline 22 decreases do~n to a minimum value at some axial loca-tion aft of the wing leading edge, generally shown at arro~ 20 in Figure 1. The presence of the engine nacelle 15 and the trailing dividing streamline near the underside of the wing 14 creates this "channel" or "nozzle" wi-th a throat at the location of arrow 20. The magnitude of the pressure reduction, and the magnitude of the loss of lift to the aircraft, is a function of nacelle position and the position lS of the fan jet dividing streamline 22 relative to the wing 1~. The more the streamline 22 "billows out" and approaches the lower wing surface, the greater the reduction in area between the wing 14 and dividing streamline 22, and thus the lower the air pressure under the wing 1~. If the position of the engine nacelle is fixed, the position of the fan jet dividing streamline 22 must be altered to decrease lift loss, allowing the aircraft -to maintain a lower angle of attack and reduce the correspond~ilg induced aerodynamic drag.
There are at least three factors that can be altered by engine designers that have an effect on the shape of the fan jet dividing streamline 22. These are the pressure of fan air exhaust, the shape of the trailing edge of the fan cowl 16, and the s.hape of the outer surface of the core cowl 18.
P~eferring no~ to Figure 3, a portion of the trailing edge of the fan cowl 16 and a portion of the core cowl 18 is shown for the purpose of explaining the influence of these three factors on the fan jet 13DV-~649 dividing streamline 22~ The space between -the aft portion of the fan cowl 16 and a forward portion of the core cowl 18 is called a bypass duct 24. The bypass duct encloses the path taken by the fan air that bypasses -the turbine section of the engine. The lines projected from the bypass duct at the aft tip of the fan cowl 16 are provided for the purpose of showing the influences of initial discharge angle, shown as 26 in Figure 3, and the exit static pressure ratio, on the shape of the dividing s-treamline 22. It can be readily appreciated from the drawing that the larger the initial discharge angle 26, the larger the maximum diameter of the dividing streamline. Similarly, the higher the exit static pressure ratio PE/Po (static pressure at exit/static pressure outside the fan cowl~, the larger the maximum diameter of the dividing streamline. The exit pressure (PE) will affect the dividing streamline because gas exiting at a higher pressure will have a greater tendency to expand radially outwardly into the surrounding airflow.
Finally, the larger the radius of the core cowl 18 relative to the engine centerline, the more the core cowl will physically force the bypass flow radially outward thereby increasing the maximum diameter of the dividing streamline. Because an increase in the maximum diameter of the dividing streamline 22 reduces the flow area between the wing lower surface and the streamline 22, pressure below the wing surface is reduced and there is an induced drag penalty as explained earlier. Any change in the construction of the fan cowl 16, bypass duct 24, and core cowl 18, which would decrease the maximum di~meter of the dividing streamline 22, will have a corresponding beneficial ef~ect on wing lift, thereby decreasing induced drag. This is the object of -the present invention.
Referring now to Figure 4, a cross-sectional view of a turbofan engine 10 is shown that incorporates one embodiment of the present invention. The invention utilizes three separate constructional features that improve the enginels bypass air exhaus-t system to reduce the maximum radius of the dividing streamline 22 and thus reduce drag. First, the trailing edge oE the inner sur~ace of the fan cowl 16 which forms the outer surface of the aft end of the bypass duct 2~, is reformed so that the downstream portion 28 of the fan cowl is curved radially inward for the purpose of directing the fan exhaust flow radially inward in respect to engine centerline. In the embodiment shown in Figure 4, the downstream portion 28 is curved radially inwardly from a position opposite the maximum radius of the core cowl 18 to the end of the bypass duct 24.
The second feature of th~ present invention is a reconstruction of the flow area distribution at the aft end of the bypass duct 24. This is accomplished by movin~ the minimum cross-sectional area or nozzle ~Jthroat ~ upstream or forward of the bypass duct exit so that the throat of the nozzle is not located where the bypass flow exhausts into the surrounding ambien-t air. By moving the nozzle throat forward, -the flow area distribution at the downstream end of -the bypass duct is increasing, thereby forming a converging-diverging nozzle. Because the bypass flow at the throat of the nozzle is choked, the bypass flow in the diverging section of the nozzle expands and looses pressure in the downstream directionO The length of the divergingsection is carefully predetermined such that the pressure at the nozzle exit is approximately equal to the ambient airstream pressure at the exit 13DV-76~9 of the fan cowl 16 during aircraft cruise operation.
This provides an exit static pressure ratio (PE/P~) of approximately 1Ø An exit static pressure ration of 1.0 causes the nozzle discharge angle of ~he bypass flow to be essentially equal to the angle of the inner wall of the fan cowl at 23. If this pressure ratio were greater than 1.0, the discharge angle would be greater than the wall angle, thus causing the jet plume to billow out relative to the b~pass wall angle.
The third feature of the present invention that reduces the maximum diameter of the dividing streamline is a reconstruction of the shape of the conic core cowl 18. ~ssentially, the conic core co~l 18 is provided with a steadily decreasing outer radius from the nozzle throat to the aft end of the core cowl. For a given amount of bypass flow passing over any core cowl at a given pressure ratio, the cowl with a lesser maximum outer radius will generally produce a lesser dividing streamline maximum diameter~
A reduction in core cowl radius provides a flow area for the fan exhaust stream that is closer to the engine 15 centerline and further from the airplane wing. This relocated flow area contributes to the effec-t of relocating the dividing streamline 22 further from the airplane wing 14.
Referring now to Figure 5, the engine and nacelle employing the present invention from Figure ~
is superimposed in dashed outline 29 on the prior art engine and nacelle of Figure 3. The differences in construction of the fan cowl 16, core cowl i8, and the downstream portion of the fan cowl 28, can be readily appreciated. A region that separates the dividing streamlines of the two engines is additionally shown with a cross-hatched section 30. The outer perimeter 32 of this cross-hatched section is the dividing streamline location for an englne employing the present invention.
The difference in proximity to the airplane wing is readily apparent.
While specific embodiments have been described, it will be apparent to those skilled in -the art that various modifications thereto can be made without departing from the scope of the invention~ as recited in the appended claims. The scope of the invention, therefore, is to be derived from the following claims.
Claims (4)
1. For a gas turbine engine effective for power-ing an aircraft, an exhaust system comprising:
an annular, first flowpath member comprising a forward portion having an outer radius increasing from a forward end of said first flowpath member to an inter-mediate portion thereof having a maximum radius; said first flowpath member further including a conical aft portion comprising first and second axially adjacent conical aft portions having an outer radius decreasing from said intermediate portion along said first and second aft portions, respectively, to an aft end of said second aft portion of said first flowpath member; and an annular, second flowpath member disposed coaxially with and spaced radially outwardly from said first flowpath member for defining therebetween an exhaust nozzle, said second flowpath member including an aft portion having an inner surface radially spaced from and surrounding said forward, intermediate and first aft portions of said first flowpath member and defining therebetween converging, throat and diverging sections, respectively, of a converging-diverging nozzle portion of said exhaust nozzle, said exhaust nozzle having a nozzle exit at a downstream end thereof for exhausting gases over said second aft portion of said first flowpath member in a generally aft direction, said gases forming a dividing streamline with airflow flowing over an outer surface of said second flowpath member during aircraft cruise operation, said streamline extending in an aft direction from an aftmost end of said second flowpath member;
said throat section being disposed upstream of said nozzle exit, and said inner surface of said aft portion of said second flowpath member having a downstream portion extending aft and radially inwardly from a position opposite said intermediate portion of said first flowpath member for directing said gases radially inwardly with respect to a central engine axis;
said diverging section having a predetermined length effective for substantially matching a pressure of said gases at said nozzle exit with a pressure of said airflow; and said converging-diverging nozzle, said downstream inner surface portion of said second flowpath member and said conical aft portion of said first flowpath member being effective for causing said dividing streamline to slope radially inwardly from said aftmost end of said second flowpath member for reducing aerodynamic drag during aircraft cruise operation.
an annular, first flowpath member comprising a forward portion having an outer radius increasing from a forward end of said first flowpath member to an inter-mediate portion thereof having a maximum radius; said first flowpath member further including a conical aft portion comprising first and second axially adjacent conical aft portions having an outer radius decreasing from said intermediate portion along said first and second aft portions, respectively, to an aft end of said second aft portion of said first flowpath member; and an annular, second flowpath member disposed coaxially with and spaced radially outwardly from said first flowpath member for defining therebetween an exhaust nozzle, said second flowpath member including an aft portion having an inner surface radially spaced from and surrounding said forward, intermediate and first aft portions of said first flowpath member and defining therebetween converging, throat and diverging sections, respectively, of a converging-diverging nozzle portion of said exhaust nozzle, said exhaust nozzle having a nozzle exit at a downstream end thereof for exhausting gases over said second aft portion of said first flowpath member in a generally aft direction, said gases forming a dividing streamline with airflow flowing over an outer surface of said second flowpath member during aircraft cruise operation, said streamline extending in an aft direction from an aftmost end of said second flowpath member;
said throat section being disposed upstream of said nozzle exit, and said inner surface of said aft portion of said second flowpath member having a downstream portion extending aft and radially inwardly from a position opposite said intermediate portion of said first flowpath member for directing said gases radially inwardly with respect to a central engine axis;
said diverging section having a predetermined length effective for substantially matching a pressure of said gases at said nozzle exit with a pressure of said airflow; and said converging-diverging nozzle, said downstream inner surface portion of said second flowpath member and said conical aft portion of said first flowpath member being effective for causing said dividing streamline to slope radially inwardly from said aftmost end of said second flowpath member for reducing aerodynamic drag during aircraft cruise operation.
2. The exhaust system according to claim 1 wherein said first flowpath member comprises a core cowl disposed about a turbine section of the engine, said second flowpath member comprises a fan cowl, and said exhaust nozzle comprises a fan bypass duct effective for discharging bypass air from said gas turbine engine.
3. An improved bypass air exhaust system for an aircraft gas turbofan engine including a fan section and a turbine section comprising:
a core cowl disposed about said turbine section and comprising a forward portion having an outer radius increasing from a forward end of said core cowl to an intermediate portion thereof having a maximum radius;
said core cowl further including a conical aft portion comprising first and second axially adjacent conical aft portions having an outer radius decreasing from said intermediate portion along said first and second aft portions, respectively, to an aft end of said second aft portion of said core cowl; and a fan cowl comprising a forward portion disposed about said fan section and an aft portion having an inner surface radially spaced from and surrounding said forward, intermediate and first aft portions of said core cowl and defining therebetween converging, throat and diverging sections, respectively, of a converging-diverging nozzle portion of a bypass duct having a duct exit at a down-stream end thereof for exhausting fan bypass air over said second aft portion of said core cowl in a generally aft direction, said fan bypass air forming a dividing stream-line with airflow flowing over an outer surface of said fan cowl during aircraft cruise operation, said streamline extending in an aft direction from an aftmost end of said fan cowl;
said throat section being disposed upstream of said duct exit, and said inner surface of said aft portion of said fan cowl having a downstream portion extending aft and radially inwardly from a position opposite said intermediate portion of said core cowl for directing said fan bypass air radially inwardly with respect to a central engine axis;
said diverging section having a predetermined length effective for substantially matching a pressure of said fan bypass air at said nozzle exit with a pressure of said airflow; and said converging-diverging nozzle, said downstream inner surface portion of said fan cowl and said conical aft portion of said core cowl being effective for causing said dividing streamline to slope radially inwardly from said aftmost end of said fan cowl for reducing aerodynamic drag during aircraft cruise operation.
4. For an aircraft of the type having a gas turbofan engine mounted below a wing, said engine having a fan cowl including a forward portion that radially surrounds a fan section and an aft portion radially spaced from and surrounding a forward portion of a core cowl defining therebetween a fan air bypass duct mounted around a central engine axis, said bypass duct having a duct exit Claim 4 - continued:
for exhausting fan bypass air over said core cowl in a generally aft direction, said fan bypass air forming a dividing streamline with ambient airflow flowing over an outer surface of said fan cowl during aircraft cruise operation, said streamline extending in an aft direction from an aftmost end of said fan cowl, an improved engine bypass air exhaust system comprising:
the core cowl further including a conical aft portion having an outer radius steadily decreasing from a position of maximum radius forward of said duct exit to an aft end of said core engine cowl; the aft portion of the fan cowl further including an aft inner surface at a downstream end of said bypass duct and extending to said duct exit thereof, said aft inner surface extending aft and radially inwardly from a position opposite said position of maximum radius of said core cowl;
said aft inner surface of said fan cowl and said outer surface of said core cowl defining therebetween a portion of said bypass duct including a converging-diverging nozzle extending to said duct exit and having a nozzle throat disposed upstream of said duct exit at said position of maximum radius of said core cowl, a converging section disposed upstream of said nozzle throat and a diverging section disposed downstream of said nozzle throat; and said diverging section of said bypass duct having an increasing flow area distribution and a predeter-mined length for the purpose of generally matching bypass air pressure to outside air pressure during aircraft cruise operation;
said converging-diverging nozzle, said aft inner surface of said fan cowl and said conical aft portion of said core cowl being effective for causing said dividing streamline to slope radially inwardly from said aftmost end of said fan cowl for reducing aerodynamic drag during
a core cowl disposed about said turbine section and comprising a forward portion having an outer radius increasing from a forward end of said core cowl to an intermediate portion thereof having a maximum radius;
said core cowl further including a conical aft portion comprising first and second axially adjacent conical aft portions having an outer radius decreasing from said intermediate portion along said first and second aft portions, respectively, to an aft end of said second aft portion of said core cowl; and a fan cowl comprising a forward portion disposed about said fan section and an aft portion having an inner surface radially spaced from and surrounding said forward, intermediate and first aft portions of said core cowl and defining therebetween converging, throat and diverging sections, respectively, of a converging-diverging nozzle portion of a bypass duct having a duct exit at a down-stream end thereof for exhausting fan bypass air over said second aft portion of said core cowl in a generally aft direction, said fan bypass air forming a dividing stream-line with airflow flowing over an outer surface of said fan cowl during aircraft cruise operation, said streamline extending in an aft direction from an aftmost end of said fan cowl;
said throat section being disposed upstream of said duct exit, and said inner surface of said aft portion of said fan cowl having a downstream portion extending aft and radially inwardly from a position opposite said intermediate portion of said core cowl for directing said fan bypass air radially inwardly with respect to a central engine axis;
said diverging section having a predetermined length effective for substantially matching a pressure of said fan bypass air at said nozzle exit with a pressure of said airflow; and said converging-diverging nozzle, said downstream inner surface portion of said fan cowl and said conical aft portion of said core cowl being effective for causing said dividing streamline to slope radially inwardly from said aftmost end of said fan cowl for reducing aerodynamic drag during aircraft cruise operation.
4. For an aircraft of the type having a gas turbofan engine mounted below a wing, said engine having a fan cowl including a forward portion that radially surrounds a fan section and an aft portion radially spaced from and surrounding a forward portion of a core cowl defining therebetween a fan air bypass duct mounted around a central engine axis, said bypass duct having a duct exit Claim 4 - continued:
for exhausting fan bypass air over said core cowl in a generally aft direction, said fan bypass air forming a dividing streamline with ambient airflow flowing over an outer surface of said fan cowl during aircraft cruise operation, said streamline extending in an aft direction from an aftmost end of said fan cowl, an improved engine bypass air exhaust system comprising:
the core cowl further including a conical aft portion having an outer radius steadily decreasing from a position of maximum radius forward of said duct exit to an aft end of said core engine cowl; the aft portion of the fan cowl further including an aft inner surface at a downstream end of said bypass duct and extending to said duct exit thereof, said aft inner surface extending aft and radially inwardly from a position opposite said position of maximum radius of said core cowl;
said aft inner surface of said fan cowl and said outer surface of said core cowl defining therebetween a portion of said bypass duct including a converging-diverging nozzle extending to said duct exit and having a nozzle throat disposed upstream of said duct exit at said position of maximum radius of said core cowl, a converging section disposed upstream of said nozzle throat and a diverging section disposed downstream of said nozzle throat; and said diverging section of said bypass duct having an increasing flow area distribution and a predeter-mined length for the purpose of generally matching bypass air pressure to outside air pressure during aircraft cruise operation;
said converging-diverging nozzle, said aft inner surface of said fan cowl and said conical aft portion of said core cowl being effective for causing said dividing streamline to slope radially inwardly from said aftmost end of said fan cowl for reducing aerodynamic drag during
Claim 4 continued:
aircraft cruise operation.
aircraft cruise operation.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US12677980A | 1980-03-03 | 1980-03-03 | |
| US126,779 | 1980-03-03 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1185101A true CA1185101A (en) | 1985-04-09 |
Family
ID=22426613
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000367228A Expired CA1185101A (en) | 1980-03-03 | 1980-12-19 | Drag-reducing nacelle |
Country Status (6)
| Country | Link |
|---|---|
| JP (1) | JPS56143330A (en) |
| CA (1) | CA1185101A (en) |
| DE (1) | DE3107496A1 (en) |
| FR (1) | FR2477100B1 (en) |
| GB (1) | GB2071769B (en) |
| IT (1) | IT1135607B (en) |
Families Citing this family (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| AU555526B2 (en) * | 1982-10-29 | 1986-09-25 | General Electric Company | Aircraft engine nacelle |
| FR2916737B1 (en) * | 2007-06-01 | 2010-05-28 | Airbus France | AIRCRAFT ENGINE ASSEMBLY WITH SLIDING CARGO. |
| US9181899B2 (en) | 2008-08-27 | 2015-11-10 | General Electric Company | Variable slope exhaust nozzle |
| US9810178B2 (en) | 2015-08-05 | 2017-11-07 | General Electric Company | Exhaust nozzle with non-coplanar and/or non-axisymmetric shape |
Family Cites Families (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1211192A (en) * | 1964-07-01 | 1970-11-04 | Gen Electric | Improvements in low drag exhaust nozzle and nacelle arrangement for turbofan engines |
| US3670964A (en) * | 1971-01-18 | 1972-06-20 | Gen Motors Corp | Jet nozzle |
| GB1420625A (en) * | 1972-08-10 | 1976-01-07 | Rolls Royce | Pitch varying mechanism for a variable pitch fan or propeller |
| US3896615A (en) * | 1973-02-08 | 1975-07-29 | United Aircraft Corp | Gas turbine engine for subsonic flight |
| US3881315A (en) * | 1973-03-19 | 1975-05-06 | Gen Electric | Fan duct flow deflector |
| CA1020365A (en) * | 1974-02-25 | 1977-11-08 | James E. Johnson | Modulating bypass variable cycle turbofan engine |
| DE2512082A1 (en) * | 1974-03-26 | 1975-10-09 | Rolls Royce 1971 Ltd | GAS TURBINE JET |
| GB1522558A (en) * | 1976-04-05 | 1978-08-23 | Rolls Royce | Duct linings |
-
1980
- 1980-12-19 CA CA000367228A patent/CA1185101A/en not_active Expired
-
1981
- 1981-02-24 IT IT19946/81A patent/IT1135607B/en active
- 1981-02-25 GB GB8105917A patent/GB2071769B/en not_active Expired
- 1981-02-27 DE DE19813107496 patent/DE3107496A1/en active Granted
- 1981-03-03 JP JP2945581A patent/JPS56143330A/en active Granted
- 1981-03-03 FR FR8104167A patent/FR2477100B1/en not_active Expired
Also Published As
| Publication number | Publication date |
|---|---|
| JPS56143330A (en) | 1981-11-09 |
| DE3107496A1 (en) | 1981-12-24 |
| FR2477100A1 (en) | 1981-09-04 |
| IT1135607B (en) | 1986-08-27 |
| GB2071769B (en) | 1984-08-22 |
| IT8119946A0 (en) | 1981-02-24 |
| GB2071769A (en) | 1981-09-23 |
| FR2477100B1 (en) | 1986-03-21 |
| DE3107496C2 (en) | 1989-12-07 |
| JPH0310560B2 (en) | 1991-02-13 |
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