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Publication numberUS3583417 A
Publication typeGrant
Publication dateJun 8, 1971
Filing dateOct 7, 1969
Priority dateOct 7, 1969
Publication numberUS 3583417 A, US 3583417A, US-A-3583417, US3583417 A, US3583417A
InventorsClark Larry T, Welliver Albertus D
Original AssigneeBoeing Co
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Sound suppressor for jet engine inlet
US 3583417 A
Abstract  available in
Images(4)
Previous page
Next page
Claims  available in
Description  (OCR text may contain errors)

United States Patent m13,ss3,417

[ 72] inventors Larry '1. Clark Seattle; Albertus D. Welllver, Kent, both 01, Wash. (21] Appl. No. 864,343 [22] Filed Oct. 7, 1969 [45] Patented June 8,1971 [73] Assignee The Boeing Company Seattle, Wash.

[54] SOUND SUPPRESSOR FOR JET ENGINE INLET 10 Claims, 8 Drawing Figs.

[52] US. Cl. l37/l5.1, 181/3321 [51] lnt.Cl. F0ln 1/16 [50] FleldoiSeareh l37/l5.i, 15.2; 181/3321 [56] References Cited FOREIGN PATENTS 921,127 3/1963 Great Britain 181/3321 Primary Examiner-Alan Cohan Attorneys-Glenn Orlob and Robert 8. Hughes ABSTRACT: An apparatus to suppress noise emitted in a forward direction from the compressor of an airplane jet engine, particularly for approach mode operation of the airplane. There are in the engine inlet two or more sets of vanes which extend transversely across the engine inlet at a location forward of the engine compressor. The sets of vanes can be translated axially with respect to one another. For low engine rotational speed the sets of vanes are positioned in generally the same transverse plane to provide a more restricted flow area in the inlet. This chokes flow in the inlet to cause the air in the inlet to flow at, or nearly at, sonic velocities in the region of the vanes. For higher engine speeds when greater air flow through the inlet is required, the two or more sets of vanes are translated axially with respect to one another to provide greater flow area through the inlet while maintaining the desired air velocity for noise attenuation.

PATENTED N 8197! SHEET 2 [IF 4 SOUND SUPPRESSOR FOR JET ENGINE INLET BACKGROUND OF THE INVENTION 1. Field of the Invention This invention relates to the suppression of noise emitted forwardly from the compressor section of a jet engine, particularly as applied to jet engines for aircraft.

2. Description of the Prior Art One way to attenuate the noise emitted forwardly from the compressor section of a jet engine is to control the flow of air through the engine in a manner that the air in the inlet reaches a velocity equal or near the speed of sound. To accomplish this, it is necessary to restrict the flow area through the inlet (i.e., choke the flow) when the engine is operating at a low power setting and reduced air intake, as for example, when the airplane is in its approach mode. On the other hand, for higher engine power requirements, the flow area through the inlet should not be restricted.

One approach of the prior art is to vary the configuration of the inlet wall so as to change the inlet area. Another prior art approach is to vary the shape and/or location of the center body of the inlet to accomplish the desired flow choking. The problems associated with a sonic throat inlet are to provide an apparatus which is light, compact and relatively uncomplicated, and yet to arrange the inlet so that the airflow therethrough be such that only an acceptable loss in engine performance be encountered in its various operating conditions.

SUMMARY OF THE INVENTION In the present invention there are two or more sets of vanes positioned in the engine inlet. These sets of vanes can be translated axially with respect to one another so that in a full choke position they are in substantially the same transverse plane to provide a restricted flow area in the inlet, in the unchoked position they are spaced axially from one another to provide a less restricted flow area in the inlet, and the vanes can be moved to partial interjacent positions to vary the effective flow area of the inlet.

Thus, it is an object of the present invention to provide an apparatus and method to attenuate noise in a jet engine by controlling airflow into the engine inlet while keeping within acceptable limits penalties in engine performance, weight, complexity and size of apparatus.

DESCRIPTION OF THE DRAWINGS FIG. 1 is an isometric view of a first embodiment of the present invention, with a portion of the engine inlet cut away;

FIG. 2 is a semischematic longitudinal sectional view of the apparatus of FIG. 1;

FIG. 3 is a semischematic longitudinal sectional view of a second embodiment of the present invention;

FIG. 4 is a fragmentary longitudinal sectional view of the first embodiment shown in FIG. 1;

FIG. 5 is a schematic drawing of a control apparatus of the present invention;

FIG. 6 is a graph illustrating the pressure recovery during cruise mode for varying Mach numbers at the choked area of the inlet when utilizing the present invention;

FIG. 7 is a graph similar to FIG. 6, but for the approach mode of the airplane;

FIG. 8 is a graph showing the reduction of noise from the engine inlet as a function of Mach number at the choked area of the inlet.

DESCRIPTION OF THE PREFERRED EMBODIMENTS In FIG. I there is shown the forward portion of an airplane jet engine 10 having an inlet 12 defining forwardly facing air intake opening 14 through which air flows to the engine compressor shown somewhat schematically at 16. At the longitudinal center axis of the engine 10 is the hub or center body 18 of the engine compressor 16.

Positioned in the inlet 12 at a location forward of the compressor I6 is a vane assembly 20, comprising three sets of inlet choking vanes, namely a forward set 22, an intermediate set 24, and a rear set 26. Each of the individual vane elements 28 is aerodynamically contoured to minimize drag with respect to air flowing through the inlet opening 14 in a direction generally parallel to the longitudinal center axis of the inlet 12. In the particular configuration shown herein, each of vanes 28 is of uniform cross section and extends horizontally in a straight line across the inlet. The front set 22 of the vanes 28 are positioned in parallel relationship one above the other in a plane perpendicular to the center axis of the inlet 12, and are mounted by their end edges to an annular frame 30 located in the annular space 32 between the inner and outer inlet walls 34 and 36, respectively. The rear set 26 of the vanes 28 is similarly positioned one above the other in a plane perpendicular to the center axis of the inlet 12.

To translate axially the forward and rear vane sets 22 and 26, there are provided three actuating mechanisms 40 placed symmetrically about the circumference of the inlet 12 in the space 32 between the inlet walls 34 and 36. (For convenience of illustration, only one of the actuating mechanisms is shown in FIGS. 1 and 4.) Each actuator 40 comprises a motor 42 which through a clutch 44 drives a shaft 46 mounted in for ward and rear bearings 48 and 50, respectively. The forward portion of the shaft 46 is formed with right-hand threads 52, while the rear portion of the shaft 46 is formed with left-hand threads 54.

There is a forward nut member 56 which threadedly engages the right-hand threaded portion 52 and a rear nut member 58 which threadedly engages the left-hand threaded portion 54. The nut member 56 is fixedly attached to the forward frame 30, while the rear nut 58 is fixedly attached to the rear frame 38. Thus, it can be seen that by rotating the shaft 46 of each of the actuating mechanisms 40 in a clockwise direction (as viewed from a position looking forwardly along the inlet axis), the forward and rear frames 30 and 38 are moved, respectively, rearwardly and forwardly so as to come closer to one another, and counterclockwise rotation of the shaft 46 moves the frames 30 and 38 away from one another.

To permit this translation of the vane sets 22 and 26, the inner inlet wall 34 is provided with suitable cutouts, as at 60, to accommodate the vane members 28. Sealing means, such as flexible lips, may be provided for these cutouts 60 as required. To limit the fore and aft movement of the vane sets 22 and 26, there are provided a forward and rear limit switch 62 and 64 respectively, to engage the rear nut member 58.

The control mechanism of the invention is indicated schematically in FIG. 5. While this type of control mechanism is described more fully in a pending patent application entitled, "Method and Systems for Controlling the Terminal Shock in an Aircraft Inlet," by Leon 0. Billig, Ser. No. 784,105, assigned to the assignee of the present application, for clarity in the presentation of the present invention, such a control apparatus will be described briefly herein.

There is a total pressure port 66 which opens forwardly into the main airstream passing into the inlet 12 so as to measure both static and dynamic pressure, and a static pressure port 68 facing laterally to the airstream at a location where the airstream is to be choked. A suitable arrangement is to locate these ports 66 and 68 in one of the vanes 28 of the intermediate vane sct 24, as shown in FIG. 4. Air from the total pressure port 66 is directed through a pair of pressure reducing orifices 70 and 72, and a first control signal is taken from a point between these orifices 70 and 72. A second control signal is taken directly from the static port 68, and these two control signals are fed into opposite sides of a bellows actuated control rod 74 of a servo valve 76. The pressure reducing orifices 70 and 72 are such that when the flow is choked, the static pressure at the port 68 will be equal to the pressure between the orifices 70 and 72.

Air from the total pressure port 66 or from an additional air supply is directed to the servo valve 76 to provide an actuating signal from either of two valve elements 78 and 80 of the servo valve 76. These valve elements 78 and 80 control a hydraulic valve 82 which in turn powers the aforementioned motor 42. Thus, if the control signal from the total pressure port 66 (taken from between the orifices 70 and 72) is higher than that from the static pressure port 68,- the control rod 74 will move to the right (as seen in FIG. 5) to signal the hydraulic actuator to drive the motor 42 in a direction to turn the shaft 46 in a clockwise direction to move the forward and rear vane sets 22 and 26 away from each other. On the other hand, if the pressure signal from the static pressure port 68 is higher than that derived from the total pressure port 66, the hydraulic actuator will cause the motor 42 to turn in the opposite direction so as to move the forward and rear vane sets 22 and 24 closer to one another and thus cause greater restriction of the airflow to the inlet 12.

The aforementioned limit switch 62 is connected to the air line between the orifices 72 and 74. Thus, when the rear vane set 26 has moved to its furthest forward position for maximum choking of the inlet 12, the pressure signal to the control rod 74 is vented through the valve 62 to prevent any further forward travel of the rearvane set 26. The other limit switch 64 is similarly connected to the static pressure line 68, so that when the rear vane set 26 has moved to the furtherest desired rearward position (at the minimum choke position) the switch 64 will open to drop the pressure in the line 68 and prevent any further rearward movement of the vane set 26.

To describe the operation of the present invention, let it be assumed that the airplane on which the jet engine is mounted is in the cruise mode, wherein the engine- 10 is operating ata power setting associated with a high mass flow into the inlet 12. In this condition, the three vane sets, 22,24 and 26 will bein the position shown in FIG. 1 (which is the solid line position shown in FIG. 4), wherein the forward set 22 is in its furthest forward position and the rear set 26 is in its furthest rearward position. In this arrangement, there is minimum flow blockage in the inlet 12. The control mechanism does not tend to translate the forward and rear vane sets 22 and 26 to cause choking of the inlet 12 during cruise because this system is turned oifduring cruise.

When the aircraft enters its approach mode, the inlet control system isienergized, the power setting of the engine is reduced, the compressor 16 demands .less air, and'the mass flow through the inlet 12 is reduced. The result is that the velocity through the vane sets22, 24 and 26 is reduced so that the static pressure at the port 68ris'es. This causes the control mechanism to respond (with the control rod 74 moving to the left and causing the motor 42 to turn clockwise) to move the forward vane set 22 rearwardly and the rear vane set 26 forwardly. Thus, the three vane sets 22, 24 and 26 begin to come into interjacent relationship with respect to one another, so that the flow'area between the vanes 28 of the intermediate set 24 begins to become moderately restricted by the leading edges of vanes 28 of the rear set 26 and further restricted by the trailing edges of the vanes 28 of the front set 22. As the power setting of the engine continues to decline and the compressor l6. demands less air so that the flow velocity in the inlet 12 decreases further, the control mechanism responds to move the vane sets 22 and 26 closer together to restrict the flow area at the vane set 24 further and thus maintain choked inlet airflow. I

When the engine 10 is moved to an increased power setting so that the compressor 16 demands more air, the flow through the vane sets, 22, 24 and 26 will begin to increase, causing a pressure drop at the static port 68. This in turn will cause the control mechanism to move the vane sets 22 and 26 further away from one another so that as the mass flow increases, the air velocity in the general region of the vane sets 22, 24 and 26 is maintained at the desired level.

Thus, it can be seen that by translating the vane sets 22, 24 and 26 axially so as to being them from an axially spaced to an interjacent relationship, the air velocity in the vicinity of the vane sets 22, 24 and 26 can be maintained at the desired level while permitting greater or less airflow through the engine de- I pending upon the engine requirements.

With reference to the graphs of FIGS. 6 and 7, the apparatus of the present invention was mounted in the inlet of an engine positioned at a stationary ground location. To obtain the values of the graph of FIG. 6, the choking vanes were placed in the cruise mode (as shown in H6. 1 and in the full lines of FIG. 4) and the engine was operated at various power settings. The curve of FIG. 7 was similarly obtained for the minimum area condition. it can be seen from the graphs of FIGS. 6 and 7 that when the Mach number in the region of the vanes reached 0.8, there was approximately a 2 percent loss in inlet total pressure.

The reduction in inlet noise in decibels was measured for the various inlet velocities at the area of the vanes. It was found that when the Mach number in the area of the vanes reached a value of about 0.8, the desired noise reduction (i.e., 8 decibels) was achieved. Thus, it can be seen that the apparatus of the present invention suppresses inlet noise while causing minimum penalty in engine performance.

A second embodiment of the present invention is illustrated in FIG. 3, wherein there is shown a forward set of vanes 84 which are stationary, and a rear set of vanes 86 which can translate axially. For the cruise mode the vanes 86 aremoved rearwardly, while for the-approach mode, the vanes 86 can be moved forwardly to be in interjacent relationship with the vanes 84 (as seen in broken lines of FIG. 3). The mode of operation of the second embodiment will be evident from the description of the operation of the first embodiment.

We claim:

- 1. In an air inlet for anair breathing engine wherein said inlet defines an inlet opening having a longitudinal axis along which air flows through said inlet opening, an apparatus for controlling airflow throughsaid inlet opening, said apparatus comprising:

a. a first set of vanes disposed in said inlet to occupy a predetermined area portion of the inlet opening,

b. a second set of vanes also disposed in said inlet to be positioned to occupy a second predetermined area portion of the inlet opening, andc. means to translate at least one of said sets of vanes with respect to the other between a first position wherein said sets of vanes are proximate one another in a generally interjacent relationship in saldinlet opening so as to provide a restricted flow area for said inlet opening, and a second position wherein said sets of vanes are positioned further from one another so as to be out of interjacent relationship with one another to provide a less restricted flow area for said inlet.

2. The apparatus as recited in claim 1, wherein there is a third set of vanes disposed in said inlet and movable to a position generally interjacent with at least one of the other sets of vanes.

3. The apparatus as recited in claim 2 wherein said second and third sets of vanes are disposed to translate in opposite directions with respect to one another.

4. The apparatus as recited in claim 3, wherein said first set of vanes is disposed to be positioned in a stationary location in said inlet, said second set is disposed to be moved between a position interjacent with said first set and a position forward of said first set, and said third set is disposed to be moved between a position interjacent with said first set and a position rearward of said first set.

5. The apparatus as recited in claim 1, wherein said vanes are disposed generally transversely to the longitudinal center axis of the inlet.

6. The apparatus as recited in claim 1, wherein there is control means operatively connected to said translating means, said control means being responsive to engine airflow requirements in a manner that for lower airflow requirements the sets of vanes move toward said first position and for greater airflow requirements the sets of vanes move toward said second position.

7. The apparatus as recited in claim 6, wherein said control means is such relative to said sets of vanes and said translating means that in response to varying engine airflow requirements said sets of vanes are positioned relative to one another to so restrict the inlet that air velocity therethrough is maintained at a level approximately near sonic velocity.

8. The apparatus as recited in claim I, wherein said vanes are aerodynamically. contoured with respect to flow along the longitudinal axis of said inlet.

9. The apparatus as recited in claim I, wherein there is a third set of vanes and said second and third sets are each mounted to a respective .one of two movable supports mounted in a cowl wall of said inlet, said translating means is operatively connected to said second and third sets in a manner that in one operating condition the second and third sets are moved, respectively, forwardly and rearwardly to move away from each other, and another operating condition the second and third sets are moved toward one another, and there is control means responsive to engine airflow requirements to cause said translating means to be in said one operating condition when there is greater engine airflow requirement, and to be in the other operating condition when there is less engine airflow requirement.

10. The apparatus as recited in claim I, wherein said apparatus is disposed in an inlet of a jet engine at a location forward of a compressor of said engine.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
GB921127A * Title not available
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3710889 *Apr 20, 1970Jan 16, 1973SnecmaAttenuation of noise from air or gas intake ducts, more especially in aircraft jet turbine engines
US3830335 *May 14, 1973Aug 20, 1974NasaNoise suppressor
US3908683 *Jun 3, 1974Sep 30, 1975Boeing CoTranslating multi-ring inlet for gas turbine engines
US4130173 *Jul 7, 1975Dec 19, 1978Vought CorporationApparatus and method for reducing flow disturbances in a flowing stream of compressible fluid
US4300656 *Sep 11, 1980Nov 17, 1981The United States Of America As Represented By The Administrator Of The National Aeronautics And Space AdministrationMultiple pure tone elimination strut assembly
US6439840 *Nov 30, 2000Aug 27, 2002Pratt & Whitney Canada Corp.Bypass duct fan noise reduction assembly
US8186942Dec 14, 2007May 29, 2012United Technologies CorporationNacelle assembly with turbulators
US8192147Dec 14, 2007Jun 5, 2012United Technologies CorporationNacelle assembly having inlet bleed
US8209953Dec 9, 2010Jul 3, 2012United Technologies CorporationGas turbine engine system providing simulated boundary layer thickness increase
US8282037Sep 30, 2010Oct 9, 2012United Technologies CorporationNacelle flow assembly
US8353164Jul 8, 2010Jan 15, 2013United Technologies CorporationGas turbine engine having slim-line nacelle
US8402739Jun 28, 2007Mar 26, 2013United Technologies CorporationVariable shape inlet section for a nacelle assembly of a gas turbine engine
US8408491Apr 24, 2007Apr 2, 2013United Technologies CorporationNacelle assembly having inlet airfoil for a gas turbine engine
US8596573Sep 5, 2012Dec 3, 2013United Technologies CorporationNacelle flow assembly
US8726632Jul 8, 2010May 20, 2014United Technologies CorporationGas turbine engine having slim-line nacelle
US8727267 *May 18, 2007May 20, 2014United Technologies CorporationVariable contraction ratio nacelle assembly for a gas turbine engine
US8844294Apr 9, 2014Sep 30, 2014United Technologies CorporationGas turbine engine having slim-line nacelle
Classifications
U.S. Classification137/15.1, 60/794, 181/214, 415/119
International ClassificationF02C7/045, F02C7/04
Cooperative ClassificationF02C7/045
European ClassificationF02C7/045