GB2259114A - Aircraft engine nacelle profile - Google Patents

Abstract

A gas turbine engine nacelle has an inlet with a profile determined by hilite points H1, H2, H2' and H3 located at extreme top, side and bottom locations of the nacelle with each hilite location being set at an individually predetermined distance from a plane which is perpendicular to the inlet centerline. An arc 104 joining hilite points H1, H2, H3 (or H1 H2' H3) defines the inlet profile. The arc may be circular or elliptical (figures 10A, 10B,) and hilite point H3 may be axially forward of hilite point H1 (Figures 9, 10A, 10B). <IMAGE>

Description

AIRCRAFT ENGINE NACELLE RAVING CIRCULAR ARC PROFILE CROSS-REFERENCE TO RELATED APPLICATION This Application is related to co-pending British application Serial No. q 1 Y ti (13DV-10268) filed concurrently herewith.

BACKGROUND OF THE INVENTION The present invention relates to nacelles for gas turbine engines of an aircraft and, more particularly, to nacelles which are constructed so that a side profile of the nacelle inlet exhibits a circular arc profile (CAP).

Conventional subsonic transport aircraft typically include wing mounted gas turbine engines which are mounted below the wing by using conventional pylons and are surrounded by.an annular conventional nacelle for providing an aerodynamically smooth envelope. With reference to FIG. 1, there is shown an exemplary subsonic commercial transport aircraft 10 powered by two turbofan engines with each engine being mounted to a wing on opposite sides of the plane.

Only one wing and engine are shown in FIG. 1.

Turbofan engine 12 is mounted to wing 14 by conventional pylon 16. Surrounding engine 12 is nacelle 18 which channels freestream airflow 20 into and about the engine 12. Engine 12 utilizes airflow 20 for combustion and the generation of thrust.

Illustrated in FIG. 2 is a vertical centerline sectional view of the nacelle 18 and engine 12 of FIG.

1. The engine 12 includes a conventional longitudinal engine centerline axis 22 which, during cruise operation of the aircraft 10, is disposed at an engine angle-of-attack aE, angle AE -being formed by the direction of the airflow 20 and the engine centerline axis 22. The nacelle 18 includes a generally annular forward nacelle inlet portion 24, and a conventional annular aft nacelle portion 26. The aft nacelle portion is spaced from the engine 12 to define a conventional bypass duct 28 and extends downstream from a conventional fan 30 of the engine 12.

During conventional operation, the engine powers the fan 30 which bypasses a portion of the freestream airflow 20 through the bypass duct for generating thrust for powering the aircraft 10. A portion of the airflow 20 is conventionally channeled through the engine 12 where it is mixed with fuel and undergoes combustion for generating combustion gases which are discharged from the engine 12 after powering, among other things, the fan 30.

With further reference to FIG. 2, the forward nacelle portion 24 includes an annular leading edge, or hilite, 34 which defines an upstream facing of a generally annular inlet face 36 which receives the freestream airflow 20 for channeling to the fan 30.

The airflow 20 which enters the inlet face 36 is also referred to as the capture streamtube 38 which enters the forward nacelle 24 through the inlet face 36.

Spillage airflow 40 is that portion of the freestream airflow which enters the inlet face 36 but not the fan 30 and is deflected around the forward nacelle 24.

The forward nacelle includes a throat 42. The throat 42 is defined as a flow region of minimum area and is positioned downstream from the leading edge 34.

An annular diffuser 44 extends downstream from the throat 42 to the fan 30. The throat is sized for channeling a predetermined mass flow rate of the airflow 20 through the fan 30. The diffuser 44 is disposed in flow communication with the inlet face 36, the throat 42, and the engine 12 and is sized and configured for reducing velocity of the airflow while increasing its static pressure.

FIG. 3 is a front perspective view of the nacelle of FIG. 2. With reference to FIG. 3, the forward nacelle inlet 24 further includes first and second transversely spaced apart sides 58 and 60, respectively, extending oppositely from the keel 52 to the crown 50 and radially outward from the inlet axis 46 which is discussed subsequently. The radial upper crown 50 and the radial lower keel 52 are transverse cross-sections of the forward nacelle 24 along a vertical plane extending through the centerline axis 22.

Conventional nacelle inlets are typically drooped wherein the nacelle diffuser has an inlet centerline or droop axis which is inclined relative to the engine centerline axis. This inlet centerline of a nacelle may be curved to correspond to the curvature of the airflow within the inlet. An example of such a curved inlet centerline can be found in U.S. Patent No.

4,722,357 to Wynosky which is incorporated herein by reference. This drooped axis arrangement allows the nacelle inlet face to be perpendicular to the freestream of airflow when the aircraft is in its design cruise mode of operation and results in minimizing installed drag over the nacelle. The acute angle formed by inlet centerline axis 46 and the engine centerline axis 22 is referred to as the droop angle, aD, and is a fixed geometric parameter. The angle formed by the engine centerline axis 22 and the airflow 20 is referred to as the engine angle-ofattack, a, which varies with changing aircraft modes of operation.

FIG. 4 is a schematic, transverse sectional view of an exemplary nacelle 18 and illustrates the coincident relationship of inlet centerline axis 46 and airflow 20 during the design cruise mode of operation. Notice that inlet face 36 is perpendicular to inlet centerline 46. Also, the engine angle-ofattack, aE, when the aircraft is in the design cruise mode of operation is equal to the droop angle, aD.

However, the engine angle-of-attack, AE S varies depending on the operating mode of the aircraft as illustrated in FIG. 5. FIG. 5 is an exemplary, transverse sectional view of nacelle 18 during a climb mode of operation. Comparing to FIG. 4, notice that the engine angle-of-attack, cr,, is greater during times when the aircraft is climbing than during the design cruise mode of operation. Thus, there is a range of over the various operating conditions.

Since weight and drag of an aircraft are important considerations, it is desirable that the nacelle be as small as possible and as light as possible for reducing weight and aerodynamic drag due to the freestream of air flowing through and around the nacelle. The length, diameter, and thickness of the nacelle are parameters which directly relate to weight and drag.

Typical aerodynamic performance parameters for evaluating low speed operation of the nacelle include total pressure recovery, circumferential pressure distortion, angle-of-attack capability of the nacelle without flow separation, and crosswind effects acting on the nacelle. At cruise operation of the aircraft, performance considerations include the variation of drag along the external surface of the nacelle due to changes in the engine airflow, freestream Mach number, and the incidence angle of the freestream airflow relative to the nacelle. The Mach number indicates the ratio of the speed of the nacelle as it travels through the air to the speed of sound in an air medium.

Furthermore, increasing environmental concerns over noise pollution have resulted in Government regulations which typically limit the amount of acceptable noise which may be directed to the ground during low speed, takeoff operation. Conventional nacelle inlets require acoustic treatment within the nacelle for meeting noise regulations and require relatively thick nacelle lower lips for meeting low speed, high angle-of-attack requirements for obtaining acceptable flow separation margin. Both of these requirements add weight to the nacelle and the relatively thick lower lip also increases drag.

One method proposed for reducing ground noise caused by aircraft turbine engines has been to employ nacelles having a scarfed or scooped inlet design. In the past, several types of scarfed or scooped inlet designs have been tested. These inlet designs are characterized by the bottom lip of the nacelle protruding in a forward manner relative to the upper lip of the nacelle, i.e. the lower lip extends forward of the conventional inlet plane. This is clearly appreciated by viewing FIGS. 6A, 6B, and 6C. FIG. 6A illustrates a side-view of deflector inlet 70. Inlet 70 receives airflow 20 and directs the airflow to engine 12. Deflector inlet 70 has a stair-shaped profile, with the forward boundaries of the inlet 70 being defined by upper leading edge 72 and lower leading edge 74.

FIG. 6B is a side view illustration of a scoop inlet 76 whose profile beginning at the upper leading edge 78 is a straight vertical line which curves to meet the lower extreme boundary defined by lower leading edge 80. FIG. 6C illustrates a side view of a scarf inlet 80 whose profile is characterized by a positively sloped straight line which connects the upper leading edge 82 with the lower leading edge 84.

The scarfed or scoop designs are further characterized by the inlet face not being perpendicular to the inlet centerline. It is apparent that the extended lower lip of the scarfed or scooped inlet design prevents noise from propagating toward the ground by reflecting noise in an upward direction.

However, any design specifications must not disregard the effects on engine performance.

During aerodynamic testing of nacelles having the scarfed or scooped inlet design it was demonstrated that the low-speed angle-of-attack (AOA) capability of such inlets was greatly improved. Unfortunately the aerodynamic performance of other parts of the nacelle, namely the top lip and sides, was compromised. The capability of airflow on the upper lip to remain attached was degraded at high flow, low AOA conditions (ground static conditions), for inlets of the type shown in FIG. 6C. The sides of the inlets tended to shed vortices and hence produce high fan face total pressure distortion when the profile shape was highly curved or discontinuous as in FIGS. 6A and 6B.

The reason these effects occur is that the air flow entering the inlet is altered by the forward or aft displacement of one lip relative to another. At mass flow ratios (MFR) greater than one, more mass flow of air is pulled into the inlet around the aft lips, and less around the forward lips, than would occur if the same lips were located in a plane normal to the inlet centerline as is the more usual convention. The reverse is true for MFR's less than one, where more mass flow of air spills out of the inlet around the aft lips and less around the forward lips. In general, this mass flow variation improves the aerodynamic performance of the forward lips, allowing smaller lip thicknesses, and degrades the aerodynamic performance of the aft lips, requiring larger lip thicknesses to regain lost performance.

Normally; one would be free to increase the thickness and length of these lips until performance requirements are met.

SUMMARY OF THE INVENTION Accordingly, it is a general object of the present invention to provide a nacelle with improved airflow characteristics.

The present invention provides in a first aspect a method for establishing an inlet profile for a nacelle of a gas turbine engine having a generally circular arc appearance, the nacelle having an inlet axis and upper, lower, and side lips distributed generally symmetrically about the inlet, comprising the steps of selecting a reference plane perpendicular to the inlet axis of said nacelle inlet; selecting a hilite point H1 on the upper lip of the nacelle inlet at a distance AX1 from said reference plane; selecting a pair of hilite points H2 on opposite sides of the inlet axis on respective ones of the side lips at a distance AX2 from the reference plane; selecting a hilite point H3 on the lower lip of the nacelle inlet at a distance AX3 from the reference plane; and conforming the.leading edge of the upper, lower, and side lips of the nacelle inlet to pass through the hilite points H1, H21 and H3, to establish a circular arc profile.

The present invention provides in a second aspect an inlet for a nacelle of a gas turbine engine comprising an upper lip and a lower lip; a hilite point H1 located at most forward location of said upper lip; a hilite point H3 located at a most forward location of said lower lip; a side lip located to the aft of said most forward location of said upper lip and to the aft of said most forward location of said lower lip, said side lip extending between and connecting said upper and lower lip; a hilite point H2 located on said side lip parallel to an inlet centerline of said nacelle; and said nacelle inlet exhibiting a generally circular arc profile.

A preferred embodiment of the present invention provides a nacelle having a throat which allows the upper and lower lips of the nacelle to be made as thin as desired to meet performance requirements by locating the side lips so as to exhibit a circular arc profile (CAP) when the nacelle is viewed in profile. The profile is determined by hilite points which are located on the top, side, and bottom lips. The hilite points are the most forward points along the inlet centerline for each of the top, side, and bottom lips.

Being so constructed, the inlet retains the low speed, AOA advantages of the scarf inlet while enhancing the performance of top lip. The side lip performance is degraded during crosswind but can be overcome by thicker lips. The CAP profile eliminates the highly curved profile or discontinuities of the scoop inlet thus preventing the shedding of vortices. The nacelle can be constructed so that the hilite points form a circular arc when viewed from the side.

BRIEF DESCRIPTION OF THE DRAWINGS A more complete appreciation of the drawings and many attendant advantages thereof will be readily obtained as the same becomes better understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:: FIG. 1 is an exemplary schematic representation of a subsonic transport aircraft having a wing mounted gas turbine engine; FIG. 2 is an exemplary transverse sectional view of the nacelle which is shown mounted to the wing of the aircraft in FIG. 1; FIG. 3 is an exemplary front view perspective illustration of the nacelle assembly shown in FIG. 2; FIG. 4 is an exemplary schematic transverse sectional view of a nacelle which illustrates the engine angle-of-attack, aE, during a cruise mode of operation; FIG. 5 is an exemplary schematic transverse sectional view of a nacelle which illustrates the engine angle-of-attack, a, during a climb mode of operation; FIG. 6A is an exemplary schematic side view illustration of a nacelle having a deflector inlet; FIG. 6B is an exemplary schematic side view of a nacelle having a scoop inlet;; FIG. 6C is an exemplary schematic side view of a nacelle having a scarf inlet; FIG. 7A is an isometric view of an exemplary CAP inlet according to one embodiment of the present invention; FIG. 7B is a side view of the nacelle of FIG. 7A and illustrates the circular arc profile; FIG. 7C is a front view of the exemplary CAP inlet of FIG. 7A; FIG. 7D is a bottom view of the nacelle of FIG.

7A; FIG. 8 is a schematic side illustration of the nacelle of FIG. 7 and demonstrates a circular arc profile; FIG. 9 is a schematic side illustration of a nacelle according to an embodiment of the present invention which illustrates how hilite points are positioned a predetermined distance from a reference plane; FIGS. lOA and lOB are side view schematic illustrations according to the present invention which illustrate how hilite points can be selected to vary the side profile of a nacelle; FIGS. llA-llC are graphs comparing the performance of conventional and CAP inlets at low speed, high angle-of-attack; and FIGS. 12 and 13 are graphs illustrating the performance improvements of CAP inlets during aircraft take-off and during engine-out climb, respectively.

When referring to the drawings, it is understood that like reference numerals designate identical or corresponding parts throughout the respective figures.

THE DETAILED DESCRIPTION OF THE INVENTION FIG. 7A is an isometric view of a nacelle 90 in accordance with one embodiment of the present -invention for housing a gas turbine engine (not shown). The nacelle includes an air inlet 92 which iS formed by the upper lip area 94, lower lip area 96, and side lip areas 98 and 100. The throat 102 represents a region of minimum flow area in the air inlet of the nacelle 90. At the most forward position on the upper lip is located the hilite point H1.

Located to the aft of the upper hilite point H1 on the side lip area 100 is hilite point H2, the most forward point of the side lip. A corresponding hilite point H21 is located on the side lip area 98. A line drawn between hilite point H2 and corresponding hilite point H2l would be perpendicular to the inlet axis of nacelle 90. At the most forward location of the lower lip is located hilite point H3. Hilite points H1 and H3 are both located forward of side hilite points H2 and H2'.

FIG. 7B is a side view of nacelle 90 and demonstrates how the contours of the nacelle form a circular arc profile (CAP) 104 which is defined by hilite points H1, H2, and H3. FIG. 7C is a front view of the exemplary CAP inlet area 102 and FIG. 7D is a bottom view of nacelle 90 which gives the reader a further appreciation of the spatial relationships between bottom hilite point H3 and side hilite point H2 and H2 In FIG. 8, a side view of the nacelle 90 is shown with the engine centerline 118 defining the center of the engine 12. The inlet axis or centerline 106 defines the center of the throat area 102. The inlet profile is indicated by CAP 104 connecting hilite points H1, H2, and H3 located along the inlet centerline axis 106.

The CAP inlet of the present invention has a profile which is determined by specifying the lip extensions AX relative to some reference plane that is perpendicular to the inlet centerline axis 106. In FIG. 9, a distance AX1 separates hilite point H1, a distance AXz separates hilite point H2/H2', and a distance AX3 separates hilite point H3 from the reference plane 122. Note that AX1 does not have to equal AX3 so that the engine inlet may have an asymmetrical configuration. A value of AX is specified for each hilite point with the three values of AX determining the circular arc profile of the nacelle. The circular arc profile gives a smooth distribution of AX from the top of the nacelle to the bottom of the nacelle unlike the discontinuous distributions inherent in prior art designs.

How far aft the hilite point H2 is located will affect the side profile of the nacelle 90 and its performance. For example, FIG. lOA is a side of the nacelle 90 and shows H1, Ha, H2ar , and H3. CAP l04B connects hilite points H1, Ha, and H3 and CAP 104A connects hilite points H1, Ha, and H3. The performance of the inlet with CAP 1048 profile will realize improvements in inlet low speed, high angle-of-attack performance compared to an inlet with a CAP 104A profile. Additionally, the profile does not have to be limited to a circular arc 104B as illustrated in FIG. lOB. FIG. 10B compares a super-elliptical profile 104C to a circular arc profile 104 B, each profile passing through the same three points.The profiles 104A, 104B, and 104C can be used to bring about different aerodynamic results for the nacelle 90.

FIGS. llA and llB illustrate how the CAP concept improves the performance of the inlet by comparing the local Mach number distribution along the bottom and side internal lips at a low speed, high AOA, power on condition. FIG. lIA shows the Mach number distribution at the keel or bottom lip for a conventional inlet and a CAP inlet. The bottom lip is the critical lip during this flight condition. Notice that the local Mach number just ahead of the shock wave (as indicated by the sudden drop in the Mach number) is -lower for the CAP inlet. The lower this Mach number, the weaker the shock becomes, the less the likelihood of flow separation, and the greater the AOA capability of the inlet. The Mach number distribution for the side lip of the CAP inlet (FIG.

llB) is overall higher than the side lip of the conventional inlet, FIG. llC. This demonstrates how the CAP concept re-distributes the lip loading to reduce the loading on the critical lip.

The aerodynamic advantages of a CAP inlet are illustrated in FIGS. 12 and 13. The data in these figures is representative of test results for the CAP concept. FIG. 12 shows the results at the low speed, high angle-of-attack, power-on condition for three freestream Mach numbers M1, M2, and M3. The increase in angle-of-attack capability as a function of CAP angle is a result of forward displacement of the bottom lip which sees less mass flow being sucked into the inlet. FIG. 13 shows the results at the low speed, moderate AOA, windmilling engine condition.

Here again, the increased AOA capability is a result of the forward displacement of'the top lip in which there is less mass flow spillage out of the inlet.

Depending on the three determinant values of the AX's, (FIG. 9) the CAP inlet of the present invention may accommodate a scarf inlet or a scoop inlet.

Furthermore, the definition of the present invention may be used to institute a circular arc profile for nacelles having an upper scoop inlet (where only the upper lip protrudes in front of the side lip) or any combination of inlet designs. The advantages of the CAP inlet concept of the present invention are improvements in performance and design flexibility.

Design flexibility arises from the CAP and its determination from the designation of the three AX's at the top, bottom, and side hilite positions of the nacelle. Also, the circular arc profile of the present invention allows an inlet to be designed for specialized improvements. For example, the upper scoop inlet improves upper lip windmill capability.

The CAP is the smoothest, lowest curvature shape that can be passed through the hilite points located at the top, side, and bottom of the nacelle, thereby avoiding the tendency to shed vortices from highly curved or discontinuous profile shapes.

The lip thickness of the sides of the nacelle 90 can be made thicker to maintain performance requirements for low speed, high flow crosswind conditions. An increase in side forebody thickness may be necessary to attach air flow under low flow, low AOA, high Mach number conditions (i.e., engineoutcruise or EROPS conditions). The flow redistribution effect caused by axial displacement of the lips, present at low speeds is also present at conditions where the Mach number may be in the 0.50.75 range. Thus, there may be a need to increase the side forebody thickness.

If the circumferential aerodynamic loading of the nacelle with CAP inlet is carefully controlled, by means of small variations of the profile shape from a pure circular arc, and by means of circumferential variation of the lip thickness and forebody thickness, it is possible to create a nacelle which has equivalent or improved performance at all relevant flow conditions.

A further advantage of the circular arc profile of the present invention results from the interaction of the nacelle flow field and the ground. Aircraft engines mounted near the ground tend to create ground vortices which can cause debris to be kicked up and ingested by the engine. Since the CAP inlet of the present invention tends to pull a greater portion of the entering air flow around the side lips during static conditions, the CAP inlet nacelle will produce a weaker ground vortex than a conventional inlet.

Thus, the foreign object damage (FOD) potential is diminished by the present invention when the engine is mounted at the same ground-to-engine centerline distance.

Furthermore, the present invention is retrofittable on conventional nacelles. As gas turbine engines and aircraft are often upgraded in size, it is frequently necessary to redesign engine inlets while preserving existing nacelle structures.

Use of this invention in redesigning an inlet may therefore save a user from the more extreme redesign efforts associated with upgrades than have been required using conventional inlet technology.

The foregoing detailed description of the preferred embodiments of the present invention is intended to be illustrative and non-limiting. Many changes and modifications are possible in light of the above teachings. Thus, it is understood that the invention may be practiced otherwise than as specifically described herein and still be within the scope of the appended claims.

Claims (14)

CLAIMS I

1. A method for establishing an inlet profile for a nacelle of a gas turbine engine having a generally circular arc appearance, the nacelle having an inlet axis and upper, lower, and side lips distributed generally symmetrically about the inlet, comprising the steps of: a) selecting a reference plane perpendicular to the inlet axis of said nacelle inlet; b) selecting a hilite point H1 on the upper lip of the nacelle inlet at a distance AX1 from said reference plane; c) selecting a pair of hilite points H on opposite sides of the inlet axis on respective ones of the side lips at a distance AX2 from the reference plane; d) selecting a hilite point H3 on the lower lip of the nacelle inlet at a distance AX3 from the reference plane; and e) conforming the leading edge of the upper, lower, and side lips of the nacelle inlet to pass through the hilite points H1, H2, and H3, to establish a circular arc profile.

2. An airstream inlet for a nacelle of a gas turbine engine, said inlet having an upper lip, and a lower lip, and a pair of opposed side lips extending between said upper and lower lips for defining a continuous, generally forward facing airstream receiving opening, said side edges having a generally arc-shaped profile with at least one of said upper lip and said lower lip extending forward of a center point of said side edges.

3. The inlet of claim 2 wherein said side profile of said side edges comprises an arc extending from said upper lip to said lower lip with each of said upper lip and said lower lip extending forward of said center point of said side edges.

4. The inlet of claim 2 wherein said arc-shaped profile begins generally at said center point of said side edges and extends to said lower lip.

5. The inlet of claim 2 wherein said arc-shaped profile comprises a partial circular arc.

6. The inlet of claim 2 wherein said arc-shaped profile comprises an elliptical arc.

7. The inlet of claim 2 wherein said side edges extend in substantially a straight line generally from said center point to said upper lip.

8. An inlet for a nacelle of a gas turbine engine comprising: an upper lip and a lower lip; a hilite point H1 located at a most forward location of said upper lip; a hilite point H3 located at a most forward location of said lower lip; a side lip located to the aft of said most forward location of said upper lip and to the aft of said most forward location of said lower lip, said side lip extending between and connecting said upper and lower lip; a hilite point H2 located on said side lip parallel to an inlet centerline of said nacelle; and said nacelle inlet exhibiting a generally circular arc profile.

9. An inlet according to claim 8 wherein said circular arc profile connects said hilite points H1 and H2 .

10. An inlet according to claim 8 wherein said circular arc profile connects said hilite points H2 and H3.

11. An inlet according to claim 8 wherein said circular arc profile connects said hilite points H H23 and H3.

12. A nacelle for a gas turbine engine having a generally circular arc profile at an inlet of the nacelle.

13. A method as claimed in Claim 1 and substantially as described with reference to the drawings.

14. An inlet as claimed in Claim 8 and substantially as described with reference to the drawings.