Document
United States Patent [191 1 1 4,220,171
Ruehr et al.
1451 Sep. 2, 1980
CURVED CENTERLINE AIR INTAKE FOR A [561 References Cited [54] GAS TURBINE ENGINE U.S. PATENT DOCUMENTS 3,011,307 12/1961 Edelfelt .
3,194,487 [75] Inventors: William C. Ruehr; James L. 7/1965 Tyler et al.
3,765,623 10/1973 Donelson .
Younghans, both of Cincinnati; 4,075,833 2/1978 Sargisson .
Edwin B. Smith, Mason, all of Ohio Primary Examiner-Alan Cohan Attorney, Agent, or Firm-Donald W. Walk; Derek P.
[73] Assignee: The United States of America as Lawrence; Norman T. Musial represented by the Administrator of the National Aeronautics and Space 1571 ABSTRACT Administration, Washington, D.C.
An inlet for a gas turbine engine is disposed about a curved centerline for the purpose of accepting intake air that is flowing at an angle to engine centerline and [21] Appl. No.: 38,980 progressively turning that intake airflow along a curved path into alignment with the engine. This curved inlet is intended for use in under-the-wing locations and similar [22] Filed: May 14, 1979 regions where airflow direction is altered by aerody- namic characteristics of the airplane. By curving the inlet, aerodynamic loss and acoustic generation and
[51] Int. C l . 3 .............................................. B64D 33/02
emission are decreased.
[52] U.S. Cl. .................................... 137/15.1; 181/214
[58] Fieldofsearch ......................... 137/1, 15.1, 15.2; 181/2 14 7 Claims, 4 Drawing Figures
. Patent Sep. 2, 1980 Sheet 1 of 2 4,220,17 1
U S . Patent Sep. 2, 1980 Sheet 2 of 2 4,220,17 1
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internal contours that redirect airflow into the engine CURVED CENTERLINE AIR INTAKE FOR A GAS without causing circumferential pressure gradients that TURBINE ENGINE propagate audible frequencies outside the engine.
It is another object of the present invention to pro- The invention herein described was made in the per- 5 vide on a gas turbine inlet with a forward intake that is formance of work under a NASA contract and is sub- angled in respect to engine centerline, an intake forward ject to the provisions of Section 305 of the National face having a minimum offset from engine centerline to Aeronautics and Space Act of 1958, Public Law 85-568 decrease aerodynamic drag caused by the nacelle inlet.
(72 Stat. 435; 42 USC 2457). These and other objects will be more fully under- 10 stood from the drawings and from the following de- BACKGROUND 0 scription, all of which are intended to be representative 1. Field of the Invention of, rather than in any way limiting on, the scope of the This invention relates to inlets for gas turbine aircraft invention.
engines. Briefly, in the apparatus of the present invention, a 2. Summary of the Prior Art 15 gas turbine engine nacelle intake is disposed about a Most current generation air intakes on subsonic com- curved centerline, a downstream end of which substan- mercial aircraft engines are drooped or angled to align tially coincides with the engine centerline for the pur- the inlet at the nacelle forward face with surrounding pose of progressively redirecting intake airflow into freestream airflow, the direction of which has been alignment with the engine centerline. The internal and odynamic characteristics of the airplane. 20 external contours of the nacelle are defined as endpoints iarly true of under-the-wing mounting of radii extending from the curved inlet centerline, and locations where the airplane wing and engine pylon the contours are generally parallel to local centerline substantially alter the direction of airflow to follow contour. With this form of construction, the inner sur- wing contour. Typically, the inlet is angled or canted face of the intake is gradually and progressively curved.
3-5 degrees to match the airflow incidence angle under 25 .The gradual curve causes intake airflow to be progres- the wing at cruise conditions. This droop or angle has sively turned in the direction of the engine centerline been achieved by orienting internal contours of the inlet thereby reducing circumferential pressure gradients symmetrically about a straight centerline, inclined to upstream of the fan and subsequent undesirable effects the engine centerline. on acoustic performance. The curved centerline inlet Testing and theoretical analysis has indicated that 30 also results in less offset of the front face of the nacelle orienting the inlet about a straight canted centerline, in as opposed to prior art straight line canted nacelles. On the manner described above, causes problems where the an engine with core-mounted gears and accessories, this engine inlet interfaces with an engine fan. In the region allows improvement of the lower nacelle external cowl- of this interface, the inner contours of the canted inlet ing and a reduction of aerodynamic drag.
must be faired into internal contours of the engine. A 35 DESCRIPTION OF THE DRAWINGS fairing at this interface abruptly turns incoming airflow from the direction of the inlet centerline to the direction While the specification concludes with the claims of the engine centerline. Tests have shown that this distinctly claiming and particularly pointing out the fairing introduces a circumferential static pressure gra- invention described herein, it is believed that the inven- dient at the interface region which increases the genera- 40 tion will be more clearly understood by reference to the tion of fan tone noise. discussion below in conjunction with the following Another feature of inlet drooping is that it affects drawings: external cowl shape and therefore nacelle drag charac- FIG. 1 is a side view of an airplane with a gas turbo- teristics. In general, the inlet external cowl shape (i.e.
fan engine mounted under the wing; projected area and length) is selected to be thin enough 45 FIG. 2 is a cross-sectional view of a prior art inlet for low wetted area and friction drag but thick enough nacelle; to avoid large spillage drag penalties. With fan cowl- FIG. 3 is a graphical comparison of static pressure as mounted gearbox and accessories, the lower nacelle a function of angular location around the nacelle wall; maximum diameter becomes much larger than the top FIG. 4 is a cross-sectional view of the present inven- and sides. Without drooping, the lower lip would be 50 tion superimposed on a phantom outline of the prior art undesirably thick. In the canted inlet, the inlet front face nacelle.
is offset downward which effectively reduces the lower DESCRIPTION OF THE PREFERRED lip thickness and is therefore desirable. However, the EMBODIMENT advent of core mounted gearbox and accessories pack- ages has eliminated the need to locally increase the 5 5 Referring now to FIG. 1, an airplane 2 is shown with lower nacelle lip thickness. Advantages are potentially a gas turbofan engine 4 mounted in an under-the-wing available because resulting decreases in projected and location. The general effect of wing and engine pylon 5 wetted areas should produce lower nacelle pressure and aerodynamics on freestream airflow is indicated by friction drag levels. While it has previously been aero- arrow 6. Essentially, the freestream airflow is deflected dynamically desirable to offset the inlet front face with 60 upward in a region just forward of the engine to follow the core-mounted gearbox and accessories with a core- an upward slant of the pylon 5 and to blend with wing mounted gearbox, it is desirable to limit the inlet offset airfoil aerodynamics. The gas turbine engine 4 is so as not to produce an unduly thin lower lip. mounted such that its inlet extends into that portion of the freestream airflow that has been deflected upwards.
SUMMARY OF THE INVENTION 65 Shown in FIG. 2 is a prior art nacelle inlet 10 that is It is, therefore, an object of the present invention to typical of current generation air intake systems for sub- provide a gas turbine inlet with a forward intake that is sonic commercial aircraft engines mounted in under- angled in respect to an engine centerline and includes the-wing locations. The nacelle inlet 10 has an inlet face
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12 that is angled along a straight centerline 13 down- use only during flight conditions, the canted inlet is not wardly so that its forward entry and centerline 13 is in normally used during static testing and, therefore, has alignment with incoming air, represented by arrow 14, been rarely used on engines undergoing static acoustic that is flowing at an angle to engine centerline 16. Be- testing.
cause freestream airflow direction is altered by engine 5 There is an additional problem that occurs during pylon and airplane wing aerodynamics, this type of inlet static testing that almost completely masks the noise nacelle offers aerodynamic advantages Of increased caused by canted inlet pressure distortion. The ad&- intake flow and reduced drag on gas turbine engines tional problem is a result of eddies in the incoming that are mounted at under-the-wing locations. A similar airflow caused by external winds, physical obstructions, type of nacelle that is canted to a lesser angle in respect 1 0 and other external factors outside the inlet nacelle.
to the engine centerline iS used O n tail-mounted engines These external factors distort the incoming airflow to compensate for alteration of freestream airflow by causing very long eddies in the flow, which, in turn, tail wing and engine pylon aerodynamics. create a major pressure distortion in the inlet that is Referring again to 2* a front fan 20* Of significantly greater than the pressure distortion caused those used on current turbofan engines, is shown di- 15 by canting the inlet. This major pressure distortion downstream Of the lo. Incoming causes tone noise that far exceeds and effectively masks airflow is directed by the nacelle 10 into this fan 20 the noise caused by the canted inlet.
where the airflow is accelerated and directed further While this masking occurs during static testing, under downstream in the engine.
flight conditions, the engine is carried at high speeds In this prior art nacelle 10, interior surface contour is 20 and the incoming airflow pattern is very different. Ex- defined around a straight inlet centerline 13. The re- ternal factors no longer eddies in the flow or any mahi% portion Of the engine “9 from the fan 2o down- resulting tone noise. Thus, the noise caused by the pres- stream, is defined around the engine centerline 16. At a sure distortion of the canted inlet becomes a significant junction between the inlet 10 and the fan 20 a relatively acoustic problem during flight conditions. Because lim- short fairing section 22 is provided to mate the straight 25 ited acoustic testing has been done under flight condi- centerline canted inlet nacelle 10 to the noncanted en- tions, and because of misconceptions regarding propa- gine and remaining portion of engine nacelle 24. At this gation of fan noise, aircraft manufacturers and those junction, large volumes of incoming airflow must be in the art are unaware Of its signifi- redirected to enter the fan 20 parallel to engine center- line 16. In the region of this junction, a flow distortion 30 cance’ in the form of a combined ,.ircumferential and radial Once it is realized that the straight centerline canted inlet is a major source of noise, the problem becomes a migration of flow occurs as the flow ap- proaches the fan face and turns to the axial direction. matter of eliminating the source of the noise and still n i S distortion in flow is characterized by ,-ircumferen- providing an inlet that directly accepts intake air that tial variation in static pressure, flow velocity, and flow 35 has been affected by engine pylon and wing aerodynm- ics. Since noise volume caused by the fan in the inlet is angle.
ProPortional to the magnitude of Pressure variation, a Circumferential distributions of inlet wall static pres- sure, just forward of the fan 20, have been measured in solution is elimination of the pressure variation that is flight on canted inlets, and the results are shown in indicated by the dashed line on the graph shown in 3.
graph form as a dashed line in FIG. 3. This dashed line 40 2* the Pressure distortion in Referring again to is a plot of local static pressure divided by average static the inlet 10 generally occurs in the region forward of pressure as a function of angular location around the circumference of the inlet nacelle 10. As the graph the fan at the fairing 22- Referring now to FIG. 4, the indicates, static pressure is lower at the bottom (180 nacelle 30 associated with the present invention is degrees) of the inlet nacelle 10 as viewed in FIG. 2 and 45 d ~ ~ n suF*mPosed on a Phantom outline ofthe Prior art nacelle 10. The present invention involves essen- higher at the top (0 degrees).
This circumferential pressure distribution is encoun- tially an inlet nacelle that is defined around a curved tered or ‘‘seen” by the fan 20 as a change in incoming centerline 32, as opposed to a straight centerline 13 for airflow velocity and flow angle at the top of the nacelle prior art nacelle 10. The coordinates of the inlet’s inter- as compared to the bottom. It is well known to those 50 nal surface 31 and external surface 33 are defined rela- that interaction of a pressure distortion tive to a large radius curve Of centerline 32 to which the skilled in the of this type with fan blades will generate tone noise. engine centerline 16 is tangent near the fan face. Center- The tone noise is produced at what is known = blade line 32 is disposed at the desired inlet flow position such passing frequencies (BPF) plus harmonics of this fre- that the inlet face 12, having a center 35, is substantially quency. Blade passing frequency caused by the distor- 55 normal to intake airflow. Inlet surface coordinates are tion pattern of FIG. 3 is equal to the number of blades defined as endpoints of radii extending from the curved times the rotational speed of the fan in revolutions per inlet centerline, and the radii extend outwardly in a direction perpendicular to local centerline contour.
second (#blades X revhec).
Interestingly, while it is well known to those skilled The curved centerline 32 allows the inlet face 12 to be in the art that a pressure distortion can interact with a 60 disposed downwardly the desired amount while elimi- rotating fan to generate tone noise, it has not been nating the necessity for the fairing 22 and its abrupt known that canted inlets cause the pressure distortion transition from inlet centerline 13 to engine centerline shown in the graph in FIG. 3 or that this pressure dis- 16. The incoming airflow is thereby progressively and tortion causes a significant amount of audible noise that gradually turned along a curved path to the direction of is emitted from the nacelle inlet. This is partly because 65 the engine centerline 16 before contacting any rotating almost all acoustic testing of aircraft engines is done on members of the fan 20. By eliminating the abrupt transi- test stands under static conditions. Since the straight tion to engine centerline, circumferential pressure dis- centerline canted inlet, shown in FIG. 2, is intended for tortion is significantly reduced or eliminated, and this
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said inlet centerline having a curved upstream por- will reduce or eliminate this source of tone noise at the tion extending from the front face center down- blade passing frequency.
stream in a curve in a point at which the engine Referring to the graph in FIG. 3, a predicted perfor- centerline is tangent to the inlet centerline; and mance result of a curved centerline inlet is shown as a said internal surface contour being curved and de- 5 solid line on the graph. The difference in pressure varia- fined as endpoints of radii extending from the tion within the two types of inlets is readily discernable, curved inlet centerline, wherein said radii extend and this has a diminishing effect on the generation of outwardly in a direction perpendicular to local noise.
centerline contour.
Besides the acoustic advantages of the curved inlet 3. The improved nacelle recited in claim 2 for a gas 10 centerline in accordance with the present invention, turbine engine in which: there are also aerodynamic advantages. Referring now the downstream end of the curved portion of the 4, it can be appreciated that there is a resulting to FIG.
nacelle inlet centerline is positioned at a point for- offset of the center of the inlet face 12 of both the ward of rotating members of the engine, and straight centerline inlet nacelle 10 and the curved cen- the nacelle inlet centerline continues from said point terline inlet nacelle 30 from engine centerline. With the in a straight line along the engine centerline.
curved centerline inlet 30, this offset of the inlet face 12 4. The apparatus recited in claim 2, and further com- is less than the corresponding offset with the straight prising: centerline inlet 10.
a generally curved forward nacelle exterior surface If the usual location of accessories and gears in the contour defined as endpoints of radii extending 2 0 lower lip 34 is changed to a core mounted location, it is outwardly from the curved inlet centerline in a possible to substantially decrease lower lip intrusion on direction perpendicular to local centerline contour.
freestream airflow and further reduce aerodynamic 5 . The apparatus recited in claim 4, and further com- drag. prising: While specific embodiments have been described, it a core-mounted gear and accessory package; and 25 will be apparent to those skilled in the art that various a lower nacelle lip having a narrow exterior outline to present a smaller front profile to freestream air- modifications thereto can be made without departing flow.
from the scope of the invention, as recited in the a p 6. A method of directing intake airflow through an pended claims. The scope of the invention, therefore, is 3 0 inlet of a nacelle into a gas turbine engine that is pro- to be derived from the following claims.
vided with an inlet face having a center displaced from ~~~i~~ described the invention, what is claimed as novel and desired to be secured by Letters Patent ofthe an engine centerline for accepting freestream airflow directly into the nacelle comprising the steps of: United States is: gradually and progressively turning the intake air- 1. An improved gas turbine intake having a front face with the flow along a curved path into with a center displaced from an engine centerline, 35 engine centerline; wherein the improvement comprises: essentially completing the alignment of intake airflow an internal surface contour based on a curved inlet prior to communication with any rotating members centerline around which inlet interior surfaces are for the purpose of eliminating circumferential pres- defined; sure gradients at and forward of the rotating mem- 40 said inlet centerline having a curved upstream por- bers.
tion extending from the front face center down-
, . A method of directing intake airflow through an
stream in a curve to a point at which the engine inlet of a nacelle into a gas turbine engine provided with centerline is tangent to the inlet centerline; and an inlet face having a center displaced from an engine said surface contour being curved and de- 45 centerline for accepting freestream airflow directly into fined as endpoints of radii extending from the the nacelle comprising the steps of: curved inlet centerline, wherein said radii extend gradually and progressively turning the intake air- outwardly in a direction perpendicular to local flow along a curved path into alignment with the centerline contour. engine centerline; 2. An improved gas turbine intake nacelle having a 50 essentially completing the alignment of intake airflow front face with a center displaced from an engine cen- prior to communication with any rotating members terline, wherein the improvement comprises: for the purpose of eliminating circumferential pres- sure gradients at and forward of the rotating mem- an internal surface contour based on a curved inlet bers, thereby reducing generation and emission of centerline around which inlet interior surfaces are noise from within the engine inlet. defined for the purpose of reducing noise genera- 55