Integrated gas turbine engine-nacelle

11 11 4,132,069

United States Patent 1191

Adamson et a l . [451 Jan. 2,1979

3,269,118 8/1966 Benedict et a l . .................... 60/39.31

INTEGRATED GAS TURBINE

1/1967 Coplin et al. ...................... 60/226 R

3,299,961 ENGINE-NACELLE

3,398,535 8/1968 Campbell et al. ................... fW39.31

Inventors: Arthur P. Adamson; Donald F. 3,442,442 5/1969 Seiwert ................................ 415/218 Sargisson, both of Cincinnati; Charles 3,485,252 12/1969 Brown ................................. 137A5.1

11/1970 Dibble et al. .......................... 4 W 7 9

3,540,682 L. Stotler, Jr., Fairfield, all of Ohio

11/1970 Johnston et al. .................... 60/39.31

3,541,794 The United States of America as Assignee:

3,542,152 11/1970 Adamson et al. ................. 60/226 R

represented by the Administrator of 3,632,460 1/1972 Palfreyman et al. ................ 415/217

3,640,357 2/1972 Etching et al. .................. 181/33 H

the National Aeronautics and Space

3,769,142 10/1973 Holmes .............................. 181/33 H

Administration, Washington, D.C.

Kasmarik et al. ................. 60/226 R

3,792,586 2/1974 Appl. No.: 753,452

3,830,058 8/1974 Ainsworth ......................... 60/226 R

3,910,374 10/1975 Holehouse ......................... 181/33 H

Filed: Dec. 22,1976

3,914,494 10/1975 Park ................................... 181/33H

Schindler ........................... 181/33 H

3,948,346 4/1976 Related U . S . Application Data Primary Examiner-Robert E. Garrett Continuation of Ser. No. 522,108,Nov. 8 , 1974, Attorney, Agent, or Firm-Robert C. Lampe, Jr.; Derek abandoned.

P. Lawrence; Norman T. Musial Int. Cl.2 .......................... FO2K 3/06, F02K 1/20; 1571 ABSTRACT F02C 3/06; F02C 7/20 U . S . Cl. ............................... 60/226 R; 60/226 A, A nacelle for use with a gas turbine engine i s provided 60/39.31; 244/54; 137A5.1; 415/201; 415/200 with an integral webbed structure resembling a spoked Field of Search .............. 60/39.31, 226 R, 226 A, wheel for rigidly interconnecting the nacelle and en- 60/262, 224; 415/79,201, 216-218, 200,209, gine. The nacelle is entirely supported in its spacial 210, 174, 191-195; 244/54; 137/15.1, 15.2; relationship with the engine by means of the webbed 181/33 H, 33 HA; 277/56 structure. The inner surface of the nacelle defines the References Cited outer limits of the engine motive fluid flow annulus while the outer surface of the nacelle defines a stream- U.S. PATENT DOCUMENTS lined envelope for the engine.

415/201 2,849,960 9/1958 Olmstead et al. ....................

277/56 3,092,393 6/1963 Morley et al. .........................

14 Claims, 5 Drawing Figures

1/1965 White .................................. 60/39.31

3,166,903

U.S. Patent ~ m . 2, 1979 Sheet 1 of 4 4,132,069

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U.S. Patent ~ m . 2,1979

Sheet 2 of 4

4,132,069

U S . Patent Jan. 2, 1979

Sheet 3 of 4 4,132,069

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U.S. Patent ~ m . 2,1979 Sheet 4 of 4 4,132,069

4,132,069 this procedure and would enable a relatively simple INTEGRATED GAS TURBINE ENGINENACELLE engine disconnect, exterior to the engine, at the pylon.

The invention herein described was made in the per- Yet another more fundamental problem has existed formance of work under a NASA contract and is sub- through non-integrationof the nacelle and engine: since ject to the provisions of Section 305 of the National 5 the responsibility for design of the various components Aeronautics and Space Act of 1958, Public Law 85-568 often lies with different manufacturers, the most aero- (72 Stat. 435; 42 USC 2457).

dynamically efficient matching of the two is not This is a continuation of application Ser. No. 522,108, achieved due to overriding individual structural consid- filed Nov. 8, 1974, now abandoned.

erations. An integrated engine-nacelle would optimize 10 engine eficiency, and thereby produce an added bonus BACKGROUND OF THE INVENTION to the performance improvement achievable through This invention relates to gas turbine engines and, the aforementioned anticipated weight reduction.

more particularly, to engine nacelles for use therewith.

The problem fachg the aircraft engine manufacturer, Jet engines for powering aircraft are provided with therefore, is to provide a lightweight nacelle integral nacelles, or other streamlined structures which envelop l5 with the engine structure which would improve overall the engine to reduce overall aerodynamic drag and performance through weight reduction and improved improve engine performance. With the advent of large- aerodynamic matching.

diameter gas turbofan engines, the required nacelle structure circumscribing the fan has become increas- SUMMARY OF THE INVENTION ingly heavy, thereby increasing aircraft weight and 2o An object of the present invention, therefore, is to reducing its range. The problem is compounded by the provide an integrated gas turbine engine nacelle which fact that since the nacelle is so large and heavy it cannot will reduce overall system weight and thereby improve be supported by the relatively lightweight, present-day aircraft performance.

gas turbine engines. It is, therefore, hung from the air- Another object of the present invention is to provide craft pylon as is the engine itself. Accordingly, there is 25 a nacelle which will eliminate structural redundancies redundancy of structure in the nacelle and engine which present in current gas turbine engine-nacelle systems.

could be eliminated with a lightweight, integrated en- Yet another object of the present invention is to pro- gine-nacelle.

vide an integrated engine-nacelle which is removable Typically, in a gas turbofan engine, a fan is provided 3o from an aircraft or other vehicle as a single unit.

forward of a core engine, the fan being rotatably driven These and other objects and advantages will be more through shaft connection by the turbine portion of the clearly understood from the following detailed descrip- engine. The fan serves to pass a large volume of air tion, the drawings and specfic examples, all of which around the core engine thereby increasing overall en- are intended to be typical of rather than in any way gine thrust. The large volume of air which bypasses the 35 limiting the scope of the present invention.

core engine (often several times the quantity of air taken Briefly stated, the above objects are attained by in- in by the core engine) is routed through an annular fan corporating lightweight compositematerials in a unique bypass duct.

structural relationship whereby the nacelle structure is The fan bypass duct is typically defined, at least in entirely supported by the engine and wherein the sup- part, by the core engine and its associated housing (or 40 porting structure is an integral structure of the gas tur- core nacelle) which comprises the inner wall of the bine engine. As a result of unitizing the engine and annulus. The outer wall is defined partially by engine nacelle structure, redundancy is eliminated. Further, structure, but predominantly by the fan nacelle which, the inner surface of the nacelle may be aerodynamically as previously noted, is supported by the pylon or air- contoured to provide the outer flow path wall of the craft wing. A shroud, or ring, is provided which cir- 45 annular fan bypass duct while the radially outward cumscribes a limited axial extent of the fan bypass duct, surface of the single thickness, though multilayered, the shroud being connected through aerodynamically nacelle serves as a streamlined envelope for the engine.

faired strut means to the core engine. This webbed The marriage of the nacelle supporting structure with structure is commonly known as the fan frame. In addi- the nacelle itself also provides greater stiffness to the tion to the aforementioned struts, a stage of guide vanes 50 assembly since the two components form one rigid is disposed across the annulus to remove any angular piece.

momentum from the flow exiting the fan to thereby increase axial thrust. The struts provide the load-carry- DESCRIPTION OF THE DRAWINGS ing structure for the shroud while the guide vanes are While the specification concludes with claims partic- loaded only in the aerodynamic sense. Integration of the 55 ularly pointing out and distinctly claiming the subject struts and guide vanes would eliminate redundancy and matter which is regarded as part of the present inven- reduce weight. The fan nacelle circumscribes the fan tion, it is believed that the invention will be more fully frame and shroud, defining the remainder of the annular understood from the following description of the pre- fan bypass flow path and, also, the outer streamlined ferred embodiment which is given in connection with envelope for the engine. Redundancy exists, therefore, 60 the accompanying drawings, in which: in both the struts and guide vanes, and in the pylon-to- FIG. 1 is a schematic representation of a gas turbofan engine and nacelle-to-pylon structure.

engine incorporating the subject invention; In addition, aircraft engine removals presently re- FIG. 2 represents an enlarged sectional view of a quire the ‘‘unbuttoning” of the nacelle in order to obtain portion of the engine of FIG. 1 , depicting in detail a access to the engine. an often awkward procedure ai - - best even when tie nacelle is of the bifurcated variety as 65 portion of the subject invention; FIG. 3 schematically depicts removal from a typical typified by US. Pat. No. 3,541,794, Johnston et al., which is assigned to the same assignee as the present aircraft pylon an engine incorporating the subject in- vention; invention. An integrated engine-nacelle would simplify 4,132,069 FIG. 4 represents a cross-sectional view of the sub- honeycomb type sandwiched between an inner wall 60 ject invention taken along line 4 4 of FIG. 1 ; and and an outer wall 62. Though not necessary, it may be FIG. 5 is similar to FIG. 4 and shows a cross-sec- desirable for acoustic purposes to perforate inner skin tional view taken along line 5-5 of FIG. 1. 60 to provide fluid communication between the motive 5 fluid passing through the engine and the cellular honey- DESCRIPTION OF THE PREFERRED comb core as is well known in the art. Further, an inter- EMBODIMENT mediate wall 64 may be provided to regulate the depth Referring to the drawings wherein like numerals of the honeycomb core 58 in communication with the correspond to l i e elements throughout, reference is motive fluid in a manner such as to “tune” the system to first directed to FIG. 1 wherein an engine depicted 10 discrete acoustic frequencies, as is also known in the art.

generally at 10 embodying the present invention is dia- Shroud 56 is also shown to include embedded therein grammatically shown. This engine may be considered an abradable insert 66 circumferentiallydisposed about as comprising generally a core engine 12, a,fan assembly the fan blade 15, which provides an abradable surface in 14 including a stage of fan blades 15, and a fan turbine the event of fan blade rub therewith. (Note that the fan 16 which is interconnected to the fan assembly 14 by 15 assembly may include blades of the variable pitch vari- shaft 18. The core engine 12 includes an axial flow ety as well as those of the fied pitch variety.) Insert 66 compressor 20 having a rotor 22. Air enters inlet 24 and may be provided with slots or grooves 68 which reduce is initially compressed by fan assembly 14. A first por- the abradable surface area (and, therefore, the rubbing tion of this compressed air enters the fan bypass duct 26 force on the blades) and which have also been shown to defined, in part, by core engine 12 and a circumscribing 20 improve fan aerodynamic performance. Additionally, a fan nacelle 28 and discharges through a fan nozzle 30. A containment ring 70 is sandwiched between the abrad- second portion of the compressed air enters inlet 32, is able insert 66 and the honeycomb core 58. Though the further compressed by the axial flow compressor 20 and preferred embodiment includes honeycomb structure as then is discharged to a combustor 34 where fuel is the core structure 58, it is within the scope of the inven- burned to provide high energy combustion gases which tion to provide a shroud of essentially solid structure drive a turbine 36. The turbine 36, in turn, drives the with or without inner and outer walls 60, 62, respec- rotor 22 through a shaft 38 in the usual manner of a gas tively.

turbine engine. The hot gases of combustion then pass A first continuous outer hoop 72 and a similar axially to and drive the fan turbine 16 which, in turn, drives the displaced second outer hoop 74 are shown to circum- fan assembly 14. A propulsive force is thus obtained by 30 scribe the core engine 12, these hoops being affied to the action of the fan assembly 14 discharging a i r from walls 60 through 64 by means of flanged braces 76 the fan bypass duct 26 through the fan nozzle 30 and by through 80 and angle braces 82 through 86. It is contem- the discharge of combustion gases from a core engine plated that outer hoops 72 and 74 could be of an integral nozzle 40 defined, in part, by plug 42. The above de- singular material thereby obviating the need for braces 35 76 through 86. Further, though only two outer hoops scription is typical of many present-day gas turbine engines and is not meant to be limiting, as it will become 72,74 are indicated, it is within the scope of the inven- readily apparent from the following description that the tion to provide one or more depending upon particular design criteria.

present invention is capable of application to any gas As is clearly shown in FIG. 2, hoops 72,74 and walls turbine enPine and is not necessarilv restricted to eas turbine eniines of the turbofan variety. The foregoLg 40 62, 64 are bonded together to form an outer annular description of the operation of the engine depicted in torque box structure while hoops 72,74 in cooperation FIG. 1 is, therefore, merely meant to be illustrative of with walls 60, 64 form a bonded inner annular torque one type of application. box structure. Braces 76,78,80,82,84 and 86 bonded to Continuing with the description of the embodiment the walls and hoops provide the necessary shear con- nections to prevent deformation of the torque boxes of the invention as depicted in FIG. 1, it is shown 45 therein that engine 10 is suspended from a pylon de- under all anticipated loadings. These torque boxes pro- picted generally at 44 which, in turn, is suspended from vide the mechanism for transferring the bending mo- 56, and the and integral with, for example, an aircraft wing 46 by ments associated with the weight of shroud aerodynamic loadings thereupon, into hoops 72,74 and means of truss structure 48. Pylon 44 is shown to in- clude, and have disposed within, aircraft accessories 50 eventually into the core engine through struts 106,108, soon to be described in greater particularity. Thus, depicted generally at 50. Critical engine accessories walls 60, 62 and 64 are load-bearing walls which pro- depicted generally at 50 are enclosed within a pocket 51 of the pylon 44 although they are an integral part of the vide circumferential support for hoops 72, 74 and the engine hardware as clearly shown in FIG. 3. Appropri- loads of which are ultimately carried by core engine 12.

ate disconnects are provided to allow separation of the 55 Intermediate wall 64 is provided to regulate the depth engine from the pylon and aircraft accessories. These of the innermost portion of honeycomb core 58 as has accessories are drivingly connected to the core engine been discussed. Thus, it is convenient to utilize this wall 12 by means of a shaft 52. Accessories access may be also as a load-bearing wall, though it is clear that in some applications the intermediate wall and the inner obtained through pylon doors 53 (FIG. 4). The engine 10 is supported on the pylon 44 by means of a thrust 60 torque box structure will not be required.

mount assembly 54, which will hereafter be described Referring now to FIGS. 2 and 4 wherein details of with more particularity. the novel nacelle configuration are more clearly de- Referring now to FIG. 2, the inventive integrated picted, it is apparent that the nacelle 28 resembles a 72 circumscribing the nacelle configuration of FIG. 1 is shown in greater wagon wheel with outer hoop core engine 12. Inner hoops 88 and 90 coaxial with detail. In particular, nacelle 28 is shown to include a 65 shroud means 56 such as a substantially cylindrical outer hoops 72, 74 are disposed within core engine 12 shroud, circumscribing fan blades 15 and a portion of and serve as means to attach the nacelle to the station- core engine 12, the shroud consisting of a core 58 of the ary core structure 92, as by bolted connections 94. Inner 4,132,069 L J hoops 88 and 90 also provide structual rigidity to the and exhaust ducts, respectively. As depicted in FIG. 1 , core engine 1 2 . Disposed radially between the outer acoustic treatment 1 3 5 , at least in part, is contemplated hoops and the inner hoops are intermediate hoops, to be of the f d depth acoustic suppression material, herein depicted as two in number, an axially forward such as honeycomb. The use of advanced composite hoop 96 and an axially rearward one 98. These hoops 5 materials is recommended such that the acoustic mate- provide structural rigidity as within a flow splitter 100 rial can be formed integrally within the duct walls and, (FIG. 2 ) which serves to split the fan motive fluid into as shown at 1 3 5 , of itself possess adequate load carrying bypass portion 26 and core portion entering inlet 3 2 . An capability. Such load bearing, full depth acoustic struc- additional pair of inner and intermediate hoops, 99 and ture, when manufactured of composite material, will 101, respectively, are provided to increase structural 10 provide significant gains in weight reduction. The inner rigidity of the flow splitter and core engine structure.

fan nacelle surface 137 (which forms the fan bypass duct Web means, such as inner struts 102 and 104, are outer flow path) and the nacelle outer surface 139 may formed integral with their respective inner and interme- be aerodynamicallycontoured to provide the most em- diate hoops and extend radially therebetween. Simi- cient codiguration.

larly, additional web means such as outer struts 106 and 15 FIG. 3 depicts schematically the entire integrated 108 are formed integral with their respective intermedi- nacelle 28 and its removal from a typical aircraft pylon ate and outer hoops, also extending radially therebe- 4 4 . Truss structure 48 includes a forward pylon mount tween. Sheaths 110 and 112 envelop the inner and outer 140 which supports the engine, in part, by pin or bolt struts 102,104 and 106,108, respectively, to provide the connection with engine hanger 142 while primary struts with aerodynamic contours. It is recognized that 20 thrust support is obtained through thrust mount assem- the sheaths may be so constructed as to provide airfoil bly 5 4 . Rear pylon mount 144 is operatively connected contours possessing characteristics such as camber and to a forward engine mount 146 by means of thrust rod stagger. As depicted in FIG. 2 , the outer struts have 148, engine mount 146 being formed integrally with been contoured to serve the function of guide vanes to intermediate hoop 98. A similar engine mount, not properly orient the motive fluid passing therethrough. 25 shown, is disposed on the other side of the engine and is The number of inner struts 102,104 and outer struts 106, connected to rear pylon mount 146 by thrust rod 150.

108 need not be equal and, in fact, the sheath outer struts Pin 152 (FIG. 2 ) facilitates connection of engine mount of FIG. 4 are depicted only schematicallysince consid- 146 and thrust rod 148. Rear pylon mount 144 is further erably more would be required to provide the solidity connected to a rear engine mount 154 (FIG. 1 ) through of a typical stage of guide vanes. In essence, a unitized, 3 0 hanger 156. Simple disconnects of a known type at wagon wheel-like nacelle frame structure is provided to pylon mounts 140, 144 permit removal of the entire entirely support the nacelle 28 upon core engine 12. In integrated engine-nacelle, with pull shaft 52 separating other words, an integrally bonded truss is formed which from the engine accessories 50 which remain within comprises the primary load-bearing structure of the pylon 44. Thus, applicants have devised a simple nacelle. It is preferred that this unitized structure be 35 method of attaching a gas turbine engine to a vehicle fabricated from lightweight, high strength composite such as an aircraft which includes the steps of first materials. Alternatively. at least a Dortion of the struc- mounting the nacelle to the enshe and then mounting ture, such as struts 102through l~ could be formed of the en&e to the vehicle. Concersely, it is foreseeabc bonded, laminated composite filaments. that the nacelle could be hung from the aircraft and the The fan nacelle 28 further includes an inlet duct 114, 40 engine then supported by the nacelle.

having a contoured lip 116 (FIG. l ) , suspended from It should be obvious to one skilled in the art that shroud 5 6 and which may be formed integral therewith certain changes can be made to the above-described or removable. A hinge connection 118 may be provided nacelle without departing from the broad, inventive to permit the inlet duct to be swung out of the way for concepts thereof. For example, in certain applications it improved access to the fan assembly 1 4 as depicted in 45 may be appropriate to support either the inlet duct 114 phantom in FIG. 3 . Alternatively, quick release fasten- or the exhaust duct 120 substantially from the pylon 44 ers of a known type could be employed, or the hinge rather than from the shroud 56. Further, in a turbojet 118 could be disposed elsewhere about the periphery of application wherein there is no fan or bypass duct, it the nacelle. Further, an exhaust duct, generally indi- would be possible to eliminate intermediate rings % , 98 cated at 120, is mounted to the axially rearward end of 50 and interconnect inner ring 88,90 and outer ring 72,74 the shroud 5 6 . FIGS. 2,3 and 5 contemplate, for exam- directly by means of integral spoked structure, the na- ple, an exhaust duct of the hinged type comprising the celle then comprising essentially a core engine nacelle.

addition of a rearward-extendingspine 122 formed inte- It i s intended that the appended claims cover these and grally with shroud 5 6 from which nacelle doors 124 and all similar variations iln applicants’ broader inventive 126 are hinged at 128 and 130, respectively. The doors 55 concepts.

are affied to shroud 56 as by cooperating tongue 132 What we claim is: and groove 134 (FIG. 2 ) formed thereupon, respec- 1 . A gas turbofan engine including a fan bypass flow 136 prevents the flow of motive fluid tively. Seal annulus having an inlet and an exit, a core engine having through the joint formed by cooperating tongue 132 an inlet duct communicating with said annulus, and a and groove 134. Access to the core engine 12 is pro- 60 nacelle comprising: vided as shown in phantom (FIG. 5). Though not pairs of axially spaced outer, intermediate and inner shown, the core engine 12 may also be provided with a coannular compositehoops, web means of compos- nacelle 138 of of the hinged or bifurcated variety. ite filament construction rigidly bonding the axially In FIG. 1 , it is anticipated that the inlet duct 114 and forwardmost of each of said pairs of hoops to- the exhaust duct 120 be constructed of lightweight, high 65 gether and further bonding the axially rearward- strength, composite material. Acoustic treatment 135 of most of each of said pairs of hoops together, the type well known in the art may be applied to the fan wherein said pair of inner hoops is disposed within bypass duct outer surface 137 formed in part by the inlet and attached to said core engine radially interior of 4,132,069

7 a

saidcoreengineinletduct,andsaidpairofinterme- 6. The gas turbofan engine as recited in claim 5 diate hoops is disposed within said core engine and wherein said duct is further characterized as an inlet extends substantially between the flow annulus and duct having a contoured lip at the inlet to said annulus.

said core engine inlet duct; 7. The gas turbofan engine as recited in claim 6 inner and outer walls of composite filament fabrica- 5 wherein said fan inlet duct is divided axially into two tion bonded to and between said outer pair of sections, one of which is bonded to one of said outer hoops to partially define said flow annulus and the hoops and the other of which is hinge connected to the outer contour of the gas turbofan engine, respec- fmed section.

tively; 8. The gas turbofan engine as recited in claim 3 fur- 10 ther comprising a load-bearing, full-depth, acoustic an intermediate wall of composite filament fabrica- tion disposed between said inner and outer walls suppression core extending radially between said inner and extending axially between said pair of outer and intermediate walls, and wherein said inner wall is hoops to which it is bonded, wherein the pair of perforated to permit flow communication between the outer hoops and the outer, intermediate and inner acoustic suppression core and the bypass flow annulus.

wallsarebondedwithshearjointstoformapairof 15 9. The gas turbofan engine as recited in claim 5 coannular inner and outer torque boxes for trans- wherein said duct is further characterized as an exhaust ferring bending moments from the walls, through duct connected to the rearwardmost of said outer hoops said web means and into said core engine; and and provided with an exhaust nozzle at the downstream composite core material bonded within said torque end thereof.

boxes. 1 0 . The gas turbofan engine as recited in claim 3 2 0 2. The gas turbofan engine as recited in claim 1 further comprising an axially extending arcuate spine wherein said web means comprises at least one strut attached to one of said outer hoops, a bifurcated duct member rigidly connecting an outer hoop to an interme- assembly split into a plurality of circumferentially ex- tending duct sectors circumscribing said core engine, diate hoop.

3. A gas turbofan engine including a fan bypass flow 25 and hinge means operatively connecting said duct sec- annulus having an inlet and an exit, a core engine having tors to said spine.

an inlet duct communicating with said annulus, and a 1 1 . The gas turbofan engine as recited in claim 4 nacelle comprising: further comprising an annular abradable insert embed- pairs of axially spaced outer and inner mannular ded within said duct substantially flush with said inner composite hoops, web means of composite filament 30 wall inner contour.

construction rigidly bonding the axially forward- 12. The gas turbofan engine as recited in claim 11 most of each of said pairs of hoops together and wherein said abradable i n s e r t includes a plurality of further bonding the axially rearwardmost of each circumferential grooves open at their radially inner of said pairs of hoops together, wherein said pair of ends and located over a stage of rotatable fan blades.

inner hoops is disposed within and attached to said 35 13. The gas turbofan engine as recited in claim 12 core engine radially interior of said core engine further comprising a high-strength containment ring within said duct and surrounding said abradable insert.

inlet duct; inner and outer walls of composite filament fabrica- 14. A gas turbofan engine including a fan bypass flow tion bonded to and between said outer pair of annulus having an inlet and an exit, a core engine having hoops to partially define said flow annulus and the 40 an inlet duct communicating with said annulus, and a outer contour of the gas turbofan engine, respec- nacelle comprising: tively; and pairs of axially spaced outer and inner coannular an intermediate wall of composite filament fabrica- composite hoops, web means of composite filament tion disposed between said inner and outer walls construction rigidly bonding the axialy forward- and extending axially between said pair of outer 45 most of each of said pairs of hoops together and hoops to which it is bonded, wherein the pair of further bonding the axially rearwardmost of each outer hoops and the outer, intermediate and inner of said pairs of hoops together, wherein said pair of walls are bonded with shear joints to form a pair of inner hoops is attached to said core engine; coannular inner and outer torque boxes for trans- inner and outer walls of composite filament fabrica- ferring bending moments from the walls, through 50 tion bonded to and between said outer pair of said web means and into said core engine. hoops to partially define said flow annulus and the outer contour of the gas turbofan engine, respec- 4. The gas turbofan engine as recited in claim 3 fur- tively; and ther comprising a generally cylindrical duct attached to wherein the pair of outer hoops and the outer and one of said outer hoops and the weight of which is inner walls are bonded with shear joints to form a transferred entirely through said torque boxes. 55 torque box for transferring bending moments from 5. The gas turbofan engine as recited in claim 4 the walls, through said web means and into said wherein said inner and outer walls extend axially from core engine. said one outer hoop to form the inner and outer con-

* * * * *

tours of said duct.