FR2477100A1 - AIR DELIVERY SYSTEM FOR TURBOCHARGED ENGINE - Google Patents

Abstract

SYSTEM ENSURING THE REQUIRED SUSTENTATION WITHOUT INCREASING THE DRAG. IT INCLUDES A DIVERGENT SECTION OF DILUTION CHANNEL 24, DOWNSTREAM OF THE TUYERE PASS, AND HAVING AN INCREASING FLOW SECTION PROFILE. APPLICATION TO ENGINES MOUNTED UNDER THE WINGS OF AN AIRPLANE.

Description

2,477,100

  The present invention relates to a nacelle and dilution channel construction for aircraft engines with

turbofan mounted under the wing.

  It is well known that an airplane wing produces lift forces during flight as a result of a

  pressure difference acting on the wing platform.

When the wing crosses an air volume, there is a relatively high air pressure under the wing and a relatively low air pressure above the wing. In general, the greater the pressure difference between the upper surface and

  the lower the wing surface, the higher the lift

  tion produced by the wing is large. It is also well known that when the angle of curvature of the wing of the airplane is more pronounced, the angle of incidence of the wing increases, thus

  that in correspondence the pressure differences and the

temptation. Unfortunately, an increase in the angle of in-

cidence also has a corresponding effect on the drag

  aerodynamics produced by the wing. Because the incident angle--

  that of the wing is increased to produce.

  temptation, the wing projects a larger frontal area

  causing increased drag.

  When an aircraft is flying at subsonic speeds, an engine under the wing of the aircraft lowers the pressures under the wing lower than they would be under the same wing without the engine. This local lowering of the pressure under the wing causes a decrease in the pressure difference and reduces the lift of the wing for a given angle of incidence. Since a given aircraft requires a fixed amount of lift to maintain its altitude at a given cruising speed, the angle of incidence must be increased to regain the amount of lift that is lost due to the presence of the gondola. engine. As expected, this increase in the angle of incidence, necessary to compensate for the loss of lift caused by the motor, causes a

additional increase in aerodynamic drag.

your commonly refer to this trail produced by the

  presence of the nacelle under the wing by "interaction trail".

  Analysis of the interaction drag revealed that different forms of engine nacelle can have in an air flow, an isolated drag similar or identical in themselves but can have very different effects on the distribution of pressures on the wing, and thus create vastly different amounts of interaction trails. Other analyzes were performed to understand these differences and the causes of this interaction trail. The results of this analysis indicate that efforts should be made to reduce the effect of the engine blower ejection system on the distribution of pressure on the wing in the

aim to reduce the interaction drag.

In short, according to an embodiment of the present invention, the construction of the nacelle and of the dilution channel of the engine is modified in order to redirect the diluted air ejected to reduce its influence on the pressures below the wing. . First, the inside profile of the inside is curved radially inwards at its rear end.

  dilution channel in order to deflect physically radial-

  inward the dilution flow and away from the underside of the wing. Second, a nozzle neck is formed in the dilution channel at a position further upstream from previous practice. The neck is located upstream in a particular place to obtain an outlet pressure of the dilution channel which closely coincides with the pressure of the outside air, so that the ejection does not relax and does not flow towards of the wing. Third, the outside diameter of a portion of the nacelle which is located immediately downstream of the outlet of the dilution channel is reduced and this portion is curved radially inward to provide a flow region for the flow.

ejection to a location further from the aircraft wing.

The rest of the description refers to the attached figures.

which represent respectively:

- Figure 1, an elevational view of a turbo engine

prior art blower mounted on a wing and its flow diagram

24771-00

  flow of the associated ejection flow, - Figure 2, a graphical representation of the local static air pressure (PS) as a function of the cross section (A) in a flow path of the nozzle or channel flow ; - Figure 3, a view of a gas turbofan engine of the prior art, partially in section and partially cut away, and the flow diagram of the dilution air flow of the engine fan; - Figure 4, a view of a gas turbofan engine

partially cut and partially cut away, which incorporates

  pore the present invention, and the flow diagram of the dilution air flow; and, - Figure 5, a sectional view of the gas turbofan engine shown in Figure 3 with superimposed in dashed lines the turbofan engine of Figure 4 which incorporates

the present invention.

Referring now to FIG. 1, there is shown a conventional turbofan engine 10, suspended by a pylon 12 under an airplane wing 14. An airplane with the engine and wing arrangement shown in FIG. 1 is designed for subsonic operation. The engine is a dilution rate turbofan aircraft engine

high standard which includes a rollover or basket 15 compre-

nant a fan cover 16 of relatively large radius at its upstream or front part and a gas generator cover 18 of relatively small radius at its downstream or rear part. The fan cover 16 covers a part of the engine fan where rotating fan blades accelerate a large volume of air towards the rear. Part of this air accelerated by the blower bypasses a part of the engine turbine and is ejected by a rear section of the blower hood 16 in the region radially surrounding the hood 18 of the gas generator. The remaining part of the blower air is sucked into the inlet 17 in the turbine part of the engine where it is used in combustion processes to produce motive energy. After passing through the turbine, the gases resulting from the combustion processes are ejected further downstream from the end

  rear 19 of the cover 18 of the gas generator.

  Analysis has shown that there are at least three main factors that influence the mutual interaction between the flow of external subsonic air adjacent to the lower surface of wing 14 and the flow of supersonic air which is discharged by the rear end of the fan cover 16. Referring again to Figure 1, a first factor is the minimum physical distance, usually indicated by the arrow

, between the lower surface of the wing and what is

  generated as the flow sharing line 22. This dividing line

flow stage is a limit between a flow of blower air ejected from the blower hood 16, and the surrounding ambient air flow which passes around the outside of the blower hood 16

  blowing. This flow sharing line is also

  naked in the art under the name of "sliding line" and is represented in its normal position during conditions

  of cruise flight by a wavy line 22.

A second factor is the total pressure ratio of the blower air flow exiting the blower hood to the ambient air pressure (PT / FAN / P0). PT / FAN

  represents the stagnation pressure of the blowing air flow

  floating ejected, and P0 the static pressure of the ambient air.

  A third factor is the Mach number of the air flow

ambient which passes outside around the bonnet 16

blooming.

  The ambient air flow between the lower surface of the wing 14 and the flow dividing line 22 is similar in certain aspects to the air flow through a channel of variable cross section. This variation in the cross section creates a "tunnel" effect on the ambient air flowing between the motor

  and the wing which is similar to the effect caused by a nozzle.

  Referring now to Figure 2, there is shown the variation of local static pressure '(P / P). In a pipe or nozzle as a function of a cross section of flow which is an approximation of the cross section between the lower surface of the wing and the flow dividing line 22 of FIG. 1. In FIG. 2, A is the local cross section,

  A :: is a reference neck or minimum section of this "conduit" -

between the wing and the engine, Ps is the local static pressure, and Pt is the stagnation pressure for a given flow. To the

  times A- and Pt are constant - for a given flow in the channel.

  The graph shows that when the flow upstream of the neck (A: :) is subsonic (M <1.0), a decrease in the section of the channel causes a decrease in the local static pressure Ps, and

when the upstream flow is supersonic (M> 1.0), an increase

tion of the channel section further causes a decrease in static pressure. This behavior is typical of the air flow in a nozzle and is well known to aeronautical engineers and mechanical engineers. The important point of this physical phenomenon is that a channel or nozzle area creates a rapid decrease in the local static pressure (PS) when the air flow changes from the subsonic state (M <1.0) to the supersonic state (M> 1.0). This is what happens between an airplane wing and an airplane engine. When the static pressure drops due to this nozzle effect in the region below the aircraft wing, it creates a detrimental effect on the

wing lift.

  Referring again to FIG. 1, the flow between the bottom surface of the wing 14 and the dividing line of the flow 22 behaves in a manner very similar to the flow crossing a channel of variable section as described below. -above. Starting from the leading edge of the wing 14, it can easily be appreciated that the distance between the bottom surface of the wing and the flow dividing line 22 decreases towards a minimum value at a certain axial location behind the leading edge of the wing, represented generally by the arrow 20 in FIG. 1. The presence of the nacelle 15 of the engine and of the leakage sharing flow line near the underside of the wing 14 create

this "channel" or "nozzle" with a neck at the location of the arrow 20.

  The magnitude of the pressure reduction and the magnitude of the aircraft lift loss is a function of the position of the nacelle and the position of the flow divider line of the blower jet 22 relative to the wing. 14. The more the flow sharing line 22. undulates as it gets closer. from the lower surface of the wing, the greater the reduction in the section between the wing 14 and the flow divider line 22, and the lower the air pressure under the wing 14. If the position of the engine nacelle is fixed we can alter the position of the flow divider of the blower jet flow to reduce the loss of lift, allowing the aircraft to maintain a

lower incidence angle and reduce aerodyna- drag

  corresponding induced mique. You can change at least three factors that have an effect on the shape of the fan jet flow divider line -10 22. These factors are the blower air ejection pressure, the shape of the trailing edge of the fan cover 16, and the shape of the outer surface of the

gas generator cover 18.

  Referring to FIG. 3, a part of the trailing edge of the fan cover 16 and a part of the cover 18 of the gas generator is shown in order to explain the influence of these three factors on the dividing line of flow of the blower jet 22. The space between the rear part of the blower cover 16 and the front part of the gas generator cover 18

  is called the dilution channel 24. The dilution channel starts

  limit the path taken by the blower air which bypasses the turbine part of the engine. The lines projected from

blower channel at the rear end of the bonnet 16

are planned to show the influences of the initial discharge angle, represented by 26 in FIG. 3, and the static outlet pressure ratio on the shape of the flow-sharing line 22. We can easily appreciate from the drawing that the larger the initial discharge angle 26, the larger the maximum diameter of the flux divide. is tall. Similarly, the higher the static pressure ratio P E / P0 (static pressure at the outlet on static pressure outside the fan cover), the larger the maximum diameter of the flow divider line. The outlet pressure (P.) will affect the flow dividing line because gases leaving at a higher pressure will have a stronger tendency to expand radially outward in the flow

surrounding air.

  Finally, the larger the radius of the cover 18 of the gas generator relative to the central axis of the engine, the more the cover of the gas generator will physically force the dilution flow radially outwards, thereby increasing the maximum diameter of the line. stream sharing. Because an increase in the maximum diameter of the flow divider line 22 reduces the flow section between the lower surface of the wing and the flow divider line: 22, the pressure under the surface of the wing is reduced and there is as explained before a

  penalizing induced drag. Any modification in the cons-

truction of the fan cover 16, of the dilution channel 24,

  and cover 18 of the gas generator, which would decrease the diameter

  maximum meter of the flow sharing line 22 would have a corresponding effect on the lift of the wing,

  thus reducing the induced drag. This is the object of the pre-

invention.

  Referring now to Figure 4, there is shown a sectional view of a turbofan engine 10 which incorporates an embodiment of the present invention. The invention uses three distinct constructive aspects which improve the ejection system of the dilution air from the engine to reduce the maximum radius of the flow divider line 22 and thus reduce the drag. First, the trailing edge of the inner surface of the fan cover 16 which forms

  the outer surface of the rear end of the di-

  lution 24, so that the downstream part 28 of the bonnet

  is curved radially inward in order to direct the blower ejection flow radially inward relative to the axis of the engine. In the embodiment shown in FIG. 4, the downstream part 28 is curved

radially inward from a position opposite

  the maximum radius of the cover 18 of the gas generator to

the end of the dilution channel 24.

The second aspect of the present invention consists in reconstructing the profile of the flow section at the rear end of the dilution channel 24. This is done by moving the minimum section or nozzle neck upstream. -. 247710t or towards the front of the outlet of the dilution channel, so that the nozzle -ol is not located where the flow of

dilutimon is ejected into the surrounding ambient air. In displacement

cutting the nozzle neck forward, we - increase the flow section at the downstream end of the

  dilution, thus forming a convergent / divergent nozzle.

  As n dilution flow at the nozzle throat is throttled, the dilution flow in the divergent part of the nozzle expands and loses pressure in the downstream direction. The

  long = r of the divergent part is carefully predetermined

  born in such a way that the pressure at the outlet of the nozzle is approximately equal to the pressure of the air flow

ambian_ at the outlet of the fan cover 16 during operation

  the aircraft in operation. This provides a static outlet pressure ratio (PE / Po) of approximately 1.0. A static outlet pressure ratio of 1.0 makes the angle of the flow

  of dilltion discharged from the nozzle is essentially equal to-

  the angle of the inner wall of the fan cover at 28.

  If this pressure ratio was greater than 1.0, the discharge angle would be greater than the angle of the wall, thus causing the plume to undulate outward relative to

  the angle of the wall of the dilution channel.

uA third aspect of the present invention which reduces the maximum diameter of the flow dividing line is a reconstruction of the shape of the conical cover 18 of the gas generator. Essentially, the conical cover 18 of the gas generator is provided with an outer radius which decreases regularly from the nozzle neck to its rear end. For a given amount of dilution flux passing over any gas generator hood at a given pressure ratio, the hood having the smallest maximum outside diameter will generally produce the smallest maximum diameter of the line.

stream sharing. A reduction in the radius of the general hood

  The gas torch provides a flow section for the fan ejector flow which is closer to the engine axis and further from the aircraft wing. This replacement of the flow section contributes to the replacement of the - 2i4-.77i1aoO -9

  * ligne- e.partage-deiflux * 22. -read farn from: l - ': - wing of -l' airplane 14.

  i do-.referring now to _figure 5, we superimposed -the mo-eur-and the nacelle.using -a present inve.ntion of the figur-_ 4 in broken lines -, 29 on the moteu-r, and the nacelle of the anterior m-rt --from lafigure 3. -, Oi can immediately appr.aier the differences in construction of the hood 16 of the blower, edu.capot 18du ga generator, .and the downstream part B- - the fan cover. There is also shown, by a hatched E = -action-, 30 a-: region which separates the partape lines from. flow of the two motors: the external perimeter 32 of this cross-hatched section is the location of the filx dividing line for the engine of the prior art, and the internal perimeter 3.4 is the location of the line of flow sharing - for = n engine using the present invention. The difference

  of r- = ximity with the wing of the aircraft appears immediately.

-.10- 2477100

Claims (4)

CLAIMS -   1. Dilution air ejection system 3eur Z - turbofan in an airplane of the type comprising a weeur à turbos = ufflante mounted under the wing, this engine having a fan cowl (16) which radially surrounds a soUffgmmtel t ^ t which has an interior surface profile at a downstream end of the blower channel which is radially curved and a blower dilution air channel (24) mounted around an axis central to the engine, this dilution channel (24) being provided with a nozzle neck placed upstream of the dm channel exit dilution (24), this dilution channel (24) being adapted for eject the fan dilution air generally towards the rear, and radially inward air relative to the center axis of the engine during operation   in cruise flight of the aircraft, system characterized in that ie -   the dilution channel (.24) is provided with a divergent section downstream from the nozzle neck with an eco-section profile growing--   2. System according to claim 1, charac * tris en what the length of the divergent section is a length-   predetermined in order to generally coincide Sa- dilution air pressure with outside air pressure   during the cruise flight of the aircraft. 3. System according to claim 1 or 2, characteris6   in that the dilution channel (24) ejects dilution air   tion to a region surrounding a conical cover (18) of gIL- gas generator, and in that the conical cover (18) of the genera-- gas meter has an external radius which decreases regularly d-min   place located in front of the exit of the dilution channel t2-4. toward a rear end of the gas generator cover.   4. Engine dilution air ejection system - turbosmuffling for an aircraft of the type having a * turbosou-flante engine mounted under the wing, this engine comprising a fan core (16) which surrounds radially a blower and a dilution channel (24) mounted around a central axis of the engine, = th dilution channel ejecting the dilution air in a general direction = rear, system characterized in that it includes means for distancing the dilution air from the sc = flower from: the wing in order to reduce the harmful effect on the lift of the wing during operation of avimn.