A variable area exit nozzle arrangement for an aircraft engine was a substantially reduced length and weight which comprises a number of longitudinally movable radial vanes and a number of fixed radial vanes. The movable radial vanes are alternately disposed with respect to the fixed radial vanes. A means is provided for displacing the movable vanes along the longitudinal axis of the engine relative to the fixed radial vanes which extend across the main exhaust flow of the engine

Latham, E. A.

Sorensen, N. E.

English

NASA Technical Reports Server

United States Patent [191 11 11 
Sorensen et al. WI 
- -  
[75] Inventors: Norman E. Sorensen, Saratoga; Eldon 
A. L.atbam, Sunnyvale, both of Calif. 
[73] Assignee: The United States of America as 
represented by the Administrator of 
the National Aeronautics and Space 
Administration, Washington, D.C. 
[21] Appl. No.: 23,436 
[22] Filed: Mar.23,1979 
[51] Int. C l . 3  .............................................. B64D 33/04 
[52] U.S. Cl. ................................... 239/127.3; 60/264; 
239/265.33 
[58] Field of Search ...................... 239/265.33, 265.19, 
239/265.11, 127.3; 60/264, 271 
[561 References Cited 
U.S. PATENT DOCUMENTS 
2,487,588 11/1949 Price ................................. 239A27.3 
3,005,308 10/1961 Bader .............................. 239/265.33 
r54i AIRCRAFT ENGINE NOZZLE 3,436,020 4/1969 Duthion et al. ................. 239/245.33 
3,642,209 2/1972 Stahl ................................ 239/265.33 
3,897,001 7/1975 Helmintolller, Jr. et al. ... 239/265.23 
Primary Examiner-Robert B. Reeves 
Asshtant Examiner-Michael J. Forman 
Attorney, Agent, or Firm-Darrell G. Brekke; John R. 
Manning; Armand McMillan 
[571 ABSTRA(X 
A variable area exit nozzle arrangement for an aircraft 
engine having a substantially reduced length and weight 
which comprises a plurality of longitudinally movable 
radial vanes and a plurality of fixed radial vanes, the 
movable radial vanes being alternately disposed with 
respect to the fixed radial vanes, and means for displac- 
ing the movable radial vanes along the longitudinal axis 
of the enghe relative to said fixed radial vanes, said 
radial vanes radially extending across the main exhaust 
flow of the engine. 
5 Claims, 7 Drawing Figures 
https://ntrs.nasa.gov/search.jsp?R=19800023884 2020-03-21T03:33:15+00:00Z
U.S. Patent Jui. 29,1980 Sheet 1 of 4 4,214,703 
I 
I U.S. Patent JU~. 29, 1980 Sheet 2 of 4 4,214,703 
U.S. Sheet 3 of 4 4,214,703 
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U.S. Patent JU~. 29, 1980 Sheet 4 of 4 4,214,703 
4,214,703 
1 2 
known plug nozzle type. The plug translates to vary the 
exit nozzle and is, of course, much longer than the noz- 
zle of the present invention and accordingly, is much 
heavier than the nozzle of the present invention. 
U.S. Pat. NO. 3,642,209, issued to Stahl, shows a noz- 
zle configuration which has little in common with the 
present invention since the outer sleeves thereof trans- 
late to vary the exit area. In addition, the nozzle, as 
shown, is only a convergent nozzle wherein the exit 
lo area is the minimum area, and thus is suitable for only 
subsonic and low supersonic applications. 
AIRCRAm ENGINE NOZZLE 
GOVERNMENT RIGHTS 
ees of the United States Government, and may be manu- 
factured and used by or for the Government for govern- 
mental purposes without payment of any royalties 
thereon or therefor. 
BACKGROUND OF THE INVENTION 
The invention herein wm made by employ-. 
1. Field of the Invention 
The present invention is directed to a short, light- 
weight jet engine nozzle that will operate efficiently 
Over the ‘Prating Mach number range Of a 
Or missile’ More the 
present invention is directed to an engine which is pro- 
vided with an arrangement for varying the area of the 
SUMMARY OF THE INVENTION 
Accordingly, it is an object of the present invention 
to provide a short, light-weight, engine nozzle that will 
operate efficiently over the complete operating Mach 
number range of a supersonic aircraft or missile. 
nozzle thereof for altering the exhaust characteristics of 
the en&e nozzle structure in accordance with opera- 20 a new and 
Another Object of the present invention is to provide 
nozzle wherein certain l’lozzle 
tion of an aircraft or -ae y t h  the subsonic and 
supersonic ranges. 
2. Description of the Prior Art 
An exemplary prior art engine-nozzle combinations is 
illustrated in FIG. 1 of the present application which 25 performance of the engine. 
shows a 1975 technology turbojet engine 8 and nozzle 9 
combination. The nozzle 9 has many movable parts, 
among which are movable tail feathers 1, inner door 
flaps 2, outer door flaps 3, secondary air Valves 4, and a 
noise suppressor 5. Movement of these elements allows 30 number of 
subsonic speeds with relatively low noise levels at take- 
elements thereof can be translated to vary the nozzle 
exit area, said translated and stationary elements form- 
ing a smooth, continuous surface at all translated posi- 
tions, thereby not adversely affecting the aerodynamic 
A further object of the present invention is to provide 
a new type of multivane, convergent-divergent nozzle, 
which is much shorter than the prior art nozzles, the 
shortness of the nozzle being directly related to the 
contained  wit^ the nozzle. 
the to Operate at both su~rsonic  and Other objects and further scope of applicability of the 
present invention will become apparent from the de- 
Off ad approach speeds* However* 
binations tailed description given hereinafter; it should be under- stood, however, that the detailed description and spe- in ’ suffer from a ‘IUUber Of disadvantages including their long length and also the 35 
heavy weight of the nozzle. cific examples, while indicating preferred embodiments 
pression is of the invention, are given by way of illustration only, 
since various changes and modifications within the 
disadvantage. spirit and scope of the invention will become apparent 
Other examples ofthe prior art include u.S. pat. NOS. 40 to those skilled in the art from this detailed description. 
3,005,308,3,897,001 and 3,642,209. All of the inventions BRIEF DESCRIPTION OF THE DRAWINGS in these patents have, in common with the present in- 
vention, translating nozzle parts to vary the nozzle exit The present invention will become more fully under- 
area. U.S. Pat. No. 3,005,308, issued to Bader, shows a stood from the detailed description given hereinbelow 
nozzle with alternate translating segments that vary the 45 and the accompanying drawings, which are given by 
area of a conventional convergent-divergent nozzle. way of illustration only, and thus are not limitative of 
The variable area nozzle arrangement of the Bader the present invention, and wherdn: 
patent Can best be Seen in FIG. 3 which Shows a plural- FIG. 1 shows an engine-nozzle combjation which is 
ity of vane segments 26 to 34, said vane segments being illustrative of the prior art, 
divided into two sets which include alternate members 50 FIG. 2 incorporates the present invention which 
thereof* A fust set Of members illustrates the much shorter, overall length of the en- 
27,29,31 and 33, which are fixed, whereas a second set to the prior 
Of =pent’ which segment members 26, 28y 309 art. Of course, such a reduction in length will also re- 
32 and 34 is axially, translatably movable with respect 
to the fixed segments. However, the nozzle disclosed in 55 FIG. 3 shows the nozzle of the present invention in its the Bader patent is much longer than that of the present 
length and light weight of the engine of the present is an view Of the Of 3; 
invention. fiso, increasing the number of translating 
segments does not change the overall length of the 60 takeoff and approach geometry; 
en@e of the Bader patent. In addition, when the seg- 
merits are translated from their “all-together” position, 
the nozzle is no longer smooth, but has many sharp 
comers and steps which lead to poor aerodynamic per- 
formance. 
U.S. Pat. No. 3,897,001, issued to Helmintoller, Jr., et 
al, merely pertains to an inlet engine nozzle system 
wherein the nozzles shown therein are of the well- 
if sound sup- 
at amff and approach speds, the 
of the suppressor 5 is also a and 
segments 
gine-nozzle com~jation when 
flect a reduction in overall weight; 
invention, and thus does not achieve the shortness in supersonic Operating position; 
FIG. 5 shows the nozzle of the present invention in its 
FIG. 6 is a cut-away perspective view of the nozzle 
Of the present invention in its supersonic design Mach 
number operating Position With all Of the vanes to- 
gether; and 
FIG. 7 is a cut-away perspective view of the nozzle 
of the present invention in its subsonic and transonic 
operating position with every other vane translated in 
the aft direction. 
65 
a 4,214,703 
3 
DETAILED DESCRIPTION O F  THE 
INVENTION 
The main advantages of the present invention are 
shown in FIG. 2 which illustrates the substantial reduc- 
tion in overall length which can be achieved by the 
present invention when compared to the prior art en- 
gine-nozzle combinations as shown in FIG. 1. As a 
matter of fact, supersonic configuration of the engine- 
nozzle combination of the present invention represents 
about a 71% reduction in length when measured from 
the turbine buckets 6 when compared to the prior art 
engines. Such a reduction in length also represents a 
substantial reduction in weight, including a greatly 
reduced overhanging weight. Also, because of its 
shorter length, the nozzle of the present invention has a 
lower radar cross-section. In addition, at take-off and 
landing speeds, a certain amount of sound suppression 
should be available from natural mixing of the external 
and internal airflows, as will be explained hereinbelow. 
Furthermore, in contrast to the prior art, the nozzle of 
the present invention only translates to provide variable 
geometry for efficient operation at both supersonic and 
subsonic speeds. 
FIG. 3 shows the nozzle of the present invention in its 
supersonic operating design position. The nozzle is 
formed mainly by tapered, radial vanes 10 and 10, for 
example 16 in number, extending from the translating 
outer sleeve 11 and the main body 12 to the translating 
inner sleeve 13 and the main inner body 14. Vanes 10 
are attached to the outer sleeve 11 and inner sleeve 13 so 
that 10,ll and 13 translate as a unit. Similarly, vanes 10 
are attached to the main body 12 and the main inner 
body 14 so that elements 10, 12 and 14 are fixed relative 
to the jet engine 15. Hot gas flow from the turbine 
buckets 6 passes through throats 16 of the nozzles 
formed by contours 17 of vanes 10 and 10. The saw- 
tooth cutouts 18, which can be 16 in number, follow the 
final Mach rhombus of the extended exit hot gas flow 
since the flow is fully expanded aft of the forward edges 
18' of the cutouts 18. The forward edges 18' of the outer 
sleeve 11 extend to eight slots 19 which allows transla- 
tion of the outer sleeve 11 without interference with the 
fixed tubes 20. Also, the forward edges 18 of the main 
body 12 extend around translating vanes 10 allowing 
free translation. Similarly, slots 21 in the main inner 
body 14 extend around vanes 10 allowing free transla- 
tion. 
Since nozzle cooling is usually required, sixteen tubes 
20 and 20 direct high-pressure cooling air from the 
forward compressor stages or from the inlet ahead of 
the engine through the hollow vanes 10 and 10' into 
chamber 26 through ports 23 communicating with the 
open ends of fixed vanes 10, through the open ends of 
translating vanes 10, and exiting through secondary 
nozzle 27 which also efficiently expands the cooling air 
flow for maximum thrust generation. Surfaces 24 can 
also further expand the main nozzle flow. Since the 
outer sleeve 11 translates in the aft direction for low- 
speed operation, every other tube 20 must telescope for 
free translation. 
FIG. 5 shows the nozzle in its take-off and approach 
geometry. Outer sleeve 11, vanes 10 and inner sleeve 13 
are translated aft by four equally spaced actuators 25 so 
that contours 17' now form the nozzle from the outer 
4 
of flow is formed in the V's 26, thus providing sound 
suppression. 
FIG. 6 is a cut-away perspective view of the nozzle 
of the present invention in its supersonic operating posi- 
5 tion with all of the vanes together, and FIG. 7 is a cut- 
away perspective view of the nozzle of the present 
invention in the subsonic and transonic operating posi- 
tion with every other vane translated in the aft direc- 
tion. By comparing FIG. 6 with FIG. 7, it can be 
readily seen that by merely translating every other vane 
of the nozzle, the throat area can be changed. 
Alternate embodiments of the present invention can 
employ any number of vanes greater than four. Each 
time the number of vanes is doubled, the nozzle length 
is halved. Thus, the present invention is concerned with 
l5 a multiple-vane, convergent-divergent nozzle which is 
much shorter than that of the prior art. Also, the trans- 
lated and stationary vanes of the present invention form 
smooth, continuous surfaces at all translated positions, 
thus contributing to an advantageous aerodynamic per- 
The invention being thus described, it will be obvious 
that the same may be varied in many ways. Such varia- 
tions are not to be regarded as a departure from the 
spirit and scope of the invention, and all such modifica- 
25 tion as would be obvious to one skilled in the art, are 
intended to be included within the scope of the follow- 
ing claims. 
2o formance. 
We claim: 
1. A variable area exit nozzle arrangement for an 
30 aircraft engine having a substantially reduced length 
and weight which comprises a plurality of longitudi- 
nally movable radial vanes and a plurality of fixed radial 
vanes, the movable radial vanes being alternately dis- 
posed with respect to the fixed radial vanes, and means 
35 for displacing the movable radial vanes along the longi- 
tudinal axis of the engine relative to said fixed radial 
vanes, said radial vanes radially extending across the 
main exhaust flow of the engine and each having a taper 
in the downstream direction whereby the exit nozzle 
40 area is varied when the movable radial vanes are dis- 
placed with respect to the fixed radial vanes, said nozzle 
including a translating outer sleeve and a translating 
inner sleeve connected by said movable radial vanes 
and a fixed main outer body and fixed main inner body 
45 connected by said fixed radial vanes, said translating 
outer and inner sleeves and movable radial vanes mov- 
ing as a unit relative to said fixed outer and inner bodies 
and said fixed radial vanes. 
2. The variable area exit nozzle arrangement of claim 
1, wherein the fixed and movable radial vanes are ta- 
50 pered from the outer sleeve and body to the inner sleeve 
and body, respectively. 
3. The variable area exit nozzle arrangement of claim 
1, wherein a plurality of fixed tubes and a plurality of 
movable tubes provide communication between said 
55 fixed radial vanes and movable radial vanes, respec- 
tively, and engine bleed air for cooling said vanes. 
4. The variable area exit nozzle arrangement of claim 
3, wherein the outer sleeve is provided with a plurality 
of longitudinally disposed slots to accommodate said 
60 plurality of movable tubes during the longitudinal trans- 
lation of said outer and inner sleeves. 
5. The variable area exit nozzle arrangement of claim 
1, wherein the fixed main inner body defines an inner 
chamber which communicates with the fixed radial 
contours of translating vanes 10. V-shaped cutouts 26 65 vanes and movable radial vanes through apertures pro- 
allow outer air flow around the nozzle to be entrained in vided in said main inner body and translating inner 
the nozzle main exhaust flow, acting much like passage- sleeve. 
way 7 shown in FIG. 1. However, now a vortex pattern * * * * *  


Aircraft engine nozzle

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