GB2071769A

GB2071769A - Drag-reducing Nacelle

Info

Publication number: GB2071769A
Application number: GB8105917A
Authority: GB
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1980-03-03
Filing date: 1981-02-25
Publication date: 1981-09-23
Also published as: DE3107496A1; JPS56143330A; IT1135607B; CA1185101A; FR2477100B1; GB2071769B; DE3107496C2; IT8119946A0; FR2477100A1; JPH0310560B2

Abstract

A wind-mounted gas turbofan engine 10 is provided with a bypass exhaust system that reduces airflow drag during subsonic aircraft flight operation by first, inwardly curving the exhaust system exit 28 to turn the bypass flow away from the airplane wing; second, by predetermining 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 core casing 18 just downstream of the bypass exhaust stream exit referenced at 29 in phantom. <IMAGE>

Description

SPECIFICATION Drag-reducing Nacelle This invention relates to nacelle and bypass duct construction for wing-mounted gas turbofan aircraft engines.

It is well known that lifting forces are produced by an aircraft wing during flight as a result of a pressure differences 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 traveling at subsonic speeds, an engine that is positioned beneath the aircraft 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 attack.

Since a given aircraft requires a fixed amount of lift to maintain altitude at a given 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 required 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 different 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.

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 diameter and curved radially inward to provide a flow region for the exhaust stream at a location that is more distant from the airplane wing.

Figure 1 is an elevation view of a prior art wing-mounted turbofan engine and its associated exhaust flow stream pattern; Figure 2 is a graphical representation of local static air pressure (Ps) as a function of crosssectional are a (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 flow 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; Figure 5 is a cross-sectional view of the gas turbofan engine shown in Figure 3 overlayed with a dashed outline of the gas turbofan engine of Figure 4 that incorporates the present invention.

Referring now to Figure 1, a conventional wing-mounted gas turbofan engine 10 is shown suspended by a pylon 12 from an airplane wing 14. 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 1 5 comprising a relatively large radius fan cowl 1 6 at its upstream or forward portion and a relatively smaller radius core cowl 1 8 at its downstream or aft region. The fan cowl 1 6 covers a fan section of the engine where rotating fan blades accelerate a large volume of air in an 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 1 6 in the region radially surrounding the core cowl 18. The remaining portion of the fan air is drawn at inlet 1 7 into the turbine section of the engine 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 14 and supersonic airflow that is discharged from the aft end of the fan cowl 16. Referring 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 respect to ambient air pressure (PT/FAN/PO). PTJFAN represents the stagnation pressure of the exhausted fan airflow, and P0 represents the ambient static air pressure.

A third factor is the Mach number of the ambient airflow which passes externally around the fan cowl 1 6.

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" effect 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 streamline 22 in Figure 1.

In explanation of Figure 2, A is the local crosssectional area, A* is a reference throat or minimum area of that "channel" between the wing and engine, P9 is local static pressure, and Pt is stagnation pressure for a given flow. Both A* and Pt are constant 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 (P,). and when the upstream flow is supersonic (M > 1.0), 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 that a channeled area or nozzle creates a rapid decrease in local static pressure (P,) 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 effect is created on wing lift.

Referring again to Figure 1, the flow between the lower surface of the wing 14 and the dividing streamline 22 behaves in a manner very similar to the flow through a duct of varying area as described above. Starting at a leading edge of the wing 14, it can be readily appreciated that the distance between the wing lower surface and the dividing streamline 22 decreases down to a minimum value at some axial location aft of the wing leading edge, generally shown at arrow 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" with 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 of the fan jet dividing streamline 22 relative to the wing 14. 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 14. 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 corresponding 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 shape of the outer surface of the core cowl 18.

Referring now to Figure 3, a portion of the trailing edge of the fan cowl 1 6 and a portion of the core cowl 1 8 is shown for the purpose of explaining the influence of these three factors on the fan jet dividing streamline 22. The space between the aft portion of the fan cowl 1 6 and a forward portion of the core cowl 1 8 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 streamline 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 PdP, (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 diameter of the dividing streamline 2?., will have a corresponding beneficial effect 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 engine's bypass air exhaust system to reduce the maximum radius of the dividing streamline 22 and thus reduce drag. First, the trailing edge of the inner surface of the fan cowl 1 6 which forms the outer surface of the aft end of the bypass duct 24, 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 the present invention is a reconstruction of the flow area distribution at the aft end of the bypass duct 24. This is accomplished by moving the minimum crosssectional area of nozzle throat 32 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 ambient air. By moving the nozzle throat forward, the flow area distribution at the downstream end of the bypass duct is increasing, thereby forming a convergingdiverging 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 direction.The length of the diverging section is carefully predetermined such that the pressure at the nozzle exit is approximately equal to the ambient airstream pressure at the exit of the fan cowl 1 6 during aircraft cruise operation. This provides an exit static pressure ratio (PE/PO) of approximately 1.0. An exit static pressure ratio of 1.0 causes the nozzle discharge angle of the bypass flow to be essentially equal to the angle of the inner wall of the fan cowl at 28. If this pressure ratio were greater than 1.9, the discharge angle would be greater than the wall angle, thus causing the jet plume to billow out relative to the bypass 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 1 8. Essentially, the conic core cowl 1 8 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's centerline and further from the airplane wing. This relocated flow area contributes to the effect 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 4 is superimposed in dashed outline 29 on the prior art engine and nacelle of Figure 3. The difference in construction of the fan cowl 16, core cowl 18, 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 crosshatched section 30. The outer perimeter 32 of this crosshatched section is the dividing streamline location for the prior art engine, and the inner perimeter 34 is the dividing streamline location for an engine employing the present invention. The difference in proximity to the airplane wing is readily apparent.

Claims

1. In an aircraft of the type having a wingmounted gas turbofan engine mounted underneath said wing, said engine having a fan cowl that radially surrounds a fan and which has an inner surface profile at a downstream end of the bypass duct that is curved radially inwardly and a fan air bypass duct mounted around a central engine axis, said bypass duct being provided with a nozzle throat positioned upstream of the bypass duct exit, said bypass duct being adapted to exhaust fan bypass air in a generally aft direction and air radially inwardly in respect to said central engine axis during aircraft cruise operation, characterized in that said bypass duct is provided with a diverging section downstream of said nozzle throat with an increasing flow area distribution.

2. The improved engine bypass air exhaust system recited in Claim 1 characterized in that the length of said diverging section is a predetermined length for the purpose of generally matching bypass air pressure to outside air pressure during aircraft cruise operation.

3. The improved engine bypass air exhaust system recited in Claim 1 or 2 characterized in that said bypass duct exhausts bypass air to a region surrounding a conic, core engine cowl, and wherein said conic, core engine cowl is provided with a steadily decreasing outer radius from a position forward of said bypass duct exit to an aft end of the core cowl.

4. In an aircraft of the type having a wingmounted gas turbofan engine mounted underneath said wing, said engine having a fan cowl that radially surrounds a fan and a fan air bypass duct mounted around a central engine axis, wherein said bypass duct exhausts fan bypass air in a generally aft direction; an improved engine bypass air exhausts system characterized by an exhaust system construction having means for directing fan bypass air away from said wing for the purpose of lessening detrimental effect on wing lift during aircraft operation.

5. A gas turbofan engine substantially as hereinbefore described with reference to and as illustrated in the drawings.