US2948115A - Drag reduction shroud for jet engines - Google Patents

Drag reduction shroud for jet engines Download PDF

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Publication number
US2948115A
US2948115A US712981A US71298158A US2948115A US 2948115 A US2948115 A US 2948115A US 712981 A US712981 A US 712981A US 71298158 A US71298158 A US 71298158A US 2948115 A US2948115 A US 2948115A
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shroud
nozzle
engine
section
openings
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US712981A
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Lorne C Dunsworth
Jr Arthur N Thomas
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Marquardt Corp
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Marquardt Corp
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K1/00Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/52Nozzles specially constructed for positioning adjacent to another nozzle or to a fixed member, e.g. fairing

Description

Aug- 9, 19 L. c. DUNSWORTH EI'AL 2,948,115

DRAG REDUCTION snaoun FOR JET ENGINES Filed Feb. 3, 1958 INVENTORS LORNE C. DUNSWORTH ARTHUR N. THOMAS JR.

rates Pater 2,948,115 Patented Aug. 9, 1960 ice DRAG REDUCTION SHROUD FOR JET ENGINES Lorne C. Dunsworth, Van Nuys, and Arthur N. Thomas,

In, N orthridge, Calif., assignors to The Marquardt Corporation, Van Nuys, Calif., a corporation of California Filed Feb. 3, 1958, Ser. No. 712,981

Claims. (Cl. 60-35.6)

This invention relates to a drag reduction shroud for jet engines and more particularly to a perforated shroud surrounding the nozzle of a jet engine to reduce drag and having openings therein to permit heat dissipation by radiation from the nozzle wall.

In high speed jet engines, such as ramjets and turbojets which are located in the air stream, it is customary to form the exit nozzle as an extension of the forward engine surface and to form the exit nozzle of thin sheet material similar to that forming the forward surface of the engine. When the complete engine is located in the airstream, the surface of the engine forward of the nozzle does not result in undue drag since the surface is substantially cylindrical. However, since the exit nozzle has a throat section and convergentand divergent sections, the nozzle surface presents a high drag contour to air flow past the engine. While the cylindrical surface of the engine could be extended to cover the exit nozzle, such an extended surface would not permit the required radiant heat dissipation from the nozzle to space in order to maintain the nozzle structure Within the temperature limit. Therefore, no covering for the exit nozzle has been utilized.

By the present invention, the drag at the exit nozzle portion of the engine is reduced while permitting suflicient radiant heat dissipation from the nozzle. A perforated cylindrical shroud is extended from the forward cylindrical surface of the engine and surrounds the exit nozzle. The shroud contains a plurality of small openings, distributed over the complete shroud or over a selected area of the shroud. Since the axis of these openings are perpendicular to the direction of air flow past the engine, they do not substantially disturb the air fiow past the shroud and thus, less drag results than when the nozzle surface is uncovered. Also, the openings are transparent to the flow of radiant energy from the nozzle to the air stream 50 that sufiicient cooling of the nozzle is provided for. Obviously, the size and the number of holes in a unit area of the shroud will be selected to provide the required heat dissipation with minimum drag. Also, the percentage of open area provided by the holes can be greater in the portion surrounding the convergent secton of the nozzle than in the portion surrounding the divergent section since the nozzle wall is subjected to higher temperatures at and forward of the throat section.

It is therefore an object of the present invention to provide a drag reduction shroud surrounding the nozzle of a jet engine and having a plurality of openings which are, transparent to flow of radiant energy from the nozzle and opaque to air flow past the engine.

Another object of the invention is to provide a cylindrical shroud for the exit nozzle of a jet engine which reduces the drag of the engine while permitting suificient heat dissipating from the nozzle.

A further object of the invention is to provide a shroud surrounding the exit nozzle *of a jet engine and containing a plurality of openings transparent to radiant energy, the percentage of open area provided by said openings being greater in the forward portion of the shroud surrounding the portion of the nozzle subjected to highest temperature.

These and other objects of the invention not specifically set forth above will become readily apparent from the accompanying description and drawings in which:

Figure 1 is a side elevational View, partly in section, of the rear portion of a jet engine and showing the plurality of openings in the shroud :which surrounds the exit nozzle.

Figure 2 is a vertical section along line 2-4. of Figure 1.

Figure 3 is an enlarged section taken along line 3-3 of Figure 2.

The embodiment of the invention chosen for illustration comprises a jet engine having a cylindrical section 5 located forward of the exit nozzle and supported by a strut 6 connected to the air frame or other structure propelled by the engine. The forward end 7 of convergent-divergent nozzle 8 is positioned within the aft end of section 5 and is attached thereto along weld line 9 in order to secure the nozzle to the engine. Nozzle 3 has a throat section It a convergent section 11 and a divergent section 12 and serves to produce thrust by discharging high temperature, high pressure gas or working fluid from the engine. A refractory coating 13 lines the interior of the cylindrical section 5 and of the nozzle in order to protectthese components from the working fiuid and reduce the temperature to which these components are subjected. A ring 14 surrounds the divergent section of nozzle 8 and is welded thereto to reinforce the nozzle structure.

A cylindrical ring 15 is partially inserted into the open end of cylindrical section 5 and is secured thereto along weld line 16. The portion of ring 15 extending aft of section 5 is received by one end of a perforated cylindrical shroud 17, which is secured to the ring by a plurality of bolts 18 spaced around the ring. The shroud 17 forms an extension of the cylindrical section 5 and completely surrounds the exit nozzle 8, thus forming space 19 between the shroud and the nozzle. Also, the aft end of the shroud extends slightly beyond the end of the nozzle.

Referring to Figures 1 and 3, the shroud 17 contains a plurality of circular openings 20 which have the same diameter and are arranged in circular rows around the shroud. The openings in each row and the rows themselves are spaced apart by the same amount and the openings in one row are staggered with respect to the openings in adjacent rows. In the area of the shroud designated A which surrounds thedivergent and throat sections of the nozzles, the spacing between the openings and the rows results in approximately fifty percent open area in the shroud. However, in the area of the shroud designated B which surrounds the divergent section of the nozzle, the spacingbetween the openings: and rows is greater so that the percentage of open areain area B ofthe shroud is considerably less than in area A. This difamount of radiant energy transfer through area A of the shroud, which surrounds the sections of the nozzle in which the gas temperature is the highest. Because of the lower gas temperature in the divergent sections 12 of the nozzle, the percentage open area in area B of the shroud can be reduced and still provide for sufficient transfer of radiant energy from the nozzle to the air stream.

During operation of the engine, hot gas or working fluid, represented by arrows C, flows through section 5 and through exit nozzle 8 to produce thrust and the air low h s he n ne s enresehted hy rro D. ince he di eot of an ow Para l l t th su eo of he shroud and perpendicular to the aggial line of, the openings, he shrou is ub a tia ly on q eto t e a r low. Als h. i no d stu bance o the art flo at the iuuetiou of se n 5 d h u 1 ihee t e sh oud s in sh h o o c o 5.- hus. h dra o h en in i elo he wh woul r su i he rr u a oz le wines e e p esent d t t e a flow r h e n However, the openings in the shroud are transparent to eeters e resen e by arrow E4, r d e r m the nozzle 3 so that heat nan be dissipated from the nozzle to the air stream. Thus, the shroud 17 results in reduceh efd n he en n a d t he ame t m permits utfi i ht too n th o The amou t o dr e c on eoo nl she by h shroud will be determined by the size of the individual qe hi i he. by reeh e e open area in h shroud resulting from the openings, and the amount of radiant energy Pa sin hrou h s ud i v hev etermin d by the percentage open area. Thus, the opening size and etse e o en. a ea ca he ar d o o t i t desired a a ce be ween d a e ti a d. ad energy trans or. F r hne sqhie rhi es. it is pr e that he. di ension o o i d u openin in he flow direction not exceed one fourth of the boundary layer thickness and that the percentage open area in the shroud he Ye hv va n he humh ro e ope i r t thanv the size of the openings. In general, the thickness of the boundary layer will increase with increase in the size ofthe engine and for a twenty-eight inch diameter engine, circular holes of quarterinch diameter have been utilized with satisfactory results. Also, it is preferable tq distribute the openings uniformly'in the shroud in order to, uniformly cool the nozzle and prevent hot spots in the nozzle. has been determined that with a fifty Per en op n. ars hrou ut th hr ud, in drag i edue d. by? abou two. nero t at e fli h speed f Mach. 2.5. With the same shroud, specificfuel.consumption is dear st-sod abou s en, ner e tat-nrorrinrum engine thrust and e l eed. y abo t h tyeroent at low cruisin thnus t; guhstantial improvement in engine. performance as. be n b ained. hr hou he. range or. th yto sixty. percent open area but, of course, the invention is not limited to this range. The density of the openings per unit area ofthev shroud can be. uniform throughout the shroud or can vary over diiferent portions, of: the shroud. as illustrated in Figure 1. Also, the. shape of the individual openings can be. other than circularand the openings can be arranged in various related patterns. ini the. shroud. Further, the shroud can be secured. to the. engine or to the nozzle in any suitable manner. Vari-. ous other modifications are contemplated by those. skilled in the. art without departing from the spirit and scope of he. nven ion, her ihette d fi d y the pp n e laim What is claimed is:

1.. In a jet engine, a cylindrical engine sectionformingan v exterion surface of said engine adjacent the air flow pastthe. engine during movement of the engine, an. exit nnz zle v secured to the; rear endof said section, and a cylindijiqalshroud secured to the. rear end of said section a portion of said-' nozzle and of such size asto be sub- 4 stantially opaque to air flow past said shroud while permitting direct flow of radiant energy from the exit nozzle to the air flow past said shroud.

2. In a jet engine as defined in claim 1 wherein said openings are distributed in said shroud to provide the largest percentage open area in the section of said shroud surrounding the portion of the nozzle of highest temperature.

3. In a jet engine, a cylindrical engine section positioned adjacent to the air flow past the engine during movement thereof, a convergent-divergent exit nozzle having its forward end inserted within the open aft end of said section, means for securing said forward end of said nozzle to said section, a mounting ring having a portion thereof inserted into said aft end of said section rearwardly of the forward end of said nozzle and having a portion extending beyond the aft end of said section, means for securing said inserted portion of said mounting ring to said section, a cylindrical shroud of substantially the same ditmeter as said; engine section and completely surrounding said nozzle, said extending portion of said mounting ring being located Within the forward end of said shroud, means for securing said forward end of said shroud to said extending portion, and a plurality of openings in said shroud located directly opposite the surface of said nozzle and of such size as to be substantially opaque to air flow past said shroud while being transparent to the direct flow of radiant energy from said exit nozzle, said openings being distributed throughout the surfac of s id. shr

4. In a jet engine as defined in claim 3 wherein said openings are distributed in said shroud to provide a larger percentage open area oppositethe converging and throat se tions of said nozzle than opposite the diverg sec tion of said nozzle.

5. In a jet engine supported within the air stream and having a straight body section forming a portion of the exterior surface of said engine an exit nozzle having its forward end connected with said body section, a shroud surrounding said exit nozzle and being ofsubstantiallythe same size and shape as said body section, said shroud containing a plurality of openings located directly opposite of at least a portion of said exit nozzle and of such size as to be substantially opaque to air flow past said shroud while permitting direct flow of radiant energy from said exit nozzle.

6. In a jet engine as defined in claim 5 wherein said openings are distributed in said shroud to provide substantially fifty percent open area in at least one portion of said shroud.

7. In a jet engine as defined in. claim 5 wherein said openings are distributed in said shroud to provide a per, centage of open area in the range between thirty and. sixty percent in at least a portion of said shroud.

8. In a jet; engine having an exit nozzle extending from a bodysection, a shroud, positioned. around said nozzle. to reduce engine. drag, and a plurality of openings in said shroud and opposite atleast a portion of said nozzle for providing passages. for the direct flow of radiantenergy from said exitnozzle, said openings being of such size and number as to be substantially opaque to airflowpast said shroud while providing for suflicient flow of radiant energy to cool said nozzle.

9. In a jet engine located at least partially in the airstream of. a movable craft, an engine section forming a,

portion of the exterior surface of said'engine and, located 5 5 opaque to air flow past said shroud While providing for References Cited in the file of this patent sutficient flow of radiant energy to cool said nozzle. UNITED STATES PATENTS 10. In a jet engine as defined in claim 9, wherein said 2 589 945 Leduc Mar 18 1952 engine section is cylindrical in shape, said shroud being of 2625O08 crook 13 1953 the same cylindrical size as said engine section to present 5 2:630:673 W011 1953 the same exterior contour to the airstream as presented by OTHER REFERENCES said engine section.

Saldin: Abstract of application Serial No. 3,040, published April 10, 1951, 645 0.6. 680-1.

US712981A 1958-02-03 1958-02-03 Drag reduction shroud for jet engines Expired - Lifetime US2948115A (en)

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Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3440820A (en) * 1967-03-13 1969-04-29 Thiokol Chemical Corp Thermal protection system for missile components subjected to excessive periods of aerodynamic heating
US3516511A (en) * 1969-05-22 1970-06-23 Rohr Corp Method and apparatus for augmenting the thrust and suppressing the noise of an aircraft jet engine
US6640537B2 (en) * 2000-12-18 2003-11-04 Pratt & Whitney Canada Corp. Aero-engine exhaust jet noise reduction assembly
US8635875B2 (en) 2010-04-29 2014-01-28 Pratt & Whitney Canada Corp. Gas turbine engine exhaust mixer including circumferentially spaced-apart radial rows of tabs extending downstream on the radial walls, crests and troughs

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2589945A (en) * 1947-02-28 1952-03-18 Leduc Rene Athodyd having air permeable converging intake section for boundary layer controls
US2625008A (en) * 1951-02-28 1953-01-13 Curtiss Wright Corp Variable flow nozzle
US2630673A (en) * 1950-09-27 1953-03-10 Gen Electric Cooling means for variable area nozzles

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2589945A (en) * 1947-02-28 1952-03-18 Leduc Rene Athodyd having air permeable converging intake section for boundary layer controls
US2630673A (en) * 1950-09-27 1953-03-10 Gen Electric Cooling means for variable area nozzles
US2625008A (en) * 1951-02-28 1953-01-13 Curtiss Wright Corp Variable flow nozzle

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3440820A (en) * 1967-03-13 1969-04-29 Thiokol Chemical Corp Thermal protection system for missile components subjected to excessive periods of aerodynamic heating
US3516511A (en) * 1969-05-22 1970-06-23 Rohr Corp Method and apparatus for augmenting the thrust and suppressing the noise of an aircraft jet engine
US6640537B2 (en) * 2000-12-18 2003-11-04 Pratt & Whitney Canada Corp. Aero-engine exhaust jet noise reduction assembly
US8635875B2 (en) 2010-04-29 2014-01-28 Pratt & Whitney Canada Corp. Gas turbine engine exhaust mixer including circumferentially spaced-apart radial rows of tabs extending downstream on the radial walls, crests and troughs

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