GB2350649A - Nozzle shroud - Google Patents

Abstract

A nozzle shroud is designed as a single-piece dismountable component to be fitted to the rear of an aircraft fuselage surrounding the propulsion engine exhaust nozzle or nozzles. The shroud is designed according to "stealth" principles so as to minimise or eliminate radar reflections. In order to reduce infra-red emissions from the hot engine exhaust the shroud is formed with at least one aperture 40, 42 through which cooler ambient air may be inducted to mix with the hot exhaust. The aperture(s) are covered with a conductive mesh or grille. The shroud is preferably constructed using lightweight material such as a composite or a honeycomb or cellular metal.

Description

NOZZLE SHROUD The invention concerns a nozzle shroud especially a shroud

for an aircraft gas turbine propulsion engine exhaust nozzle. In particular the invention relates to a dismountable nozzle shroud possessing stealthy characteristics.

In the case of certain types of aircraft it may be desirable to implement steps the objective of which is to reduce characteristics by which the aircraft may be detected and/or tracked. In particular, by detection of its radar signature and infra-red emissions caused by the propulsion engine exhaust nozzle(s). The problem of radar signature is addressed by design, rules, principles and practices which minimise, or even eliminate, radar reflections from an aircraft and to some extent at least are well known or reasonably familiar. To mention a few by way of example, the use of carefully designed surfaces in place of large flat planes, edges and joints being saw-toothed rather than straight lines, and the use of radar absorbent materials on exposed surfaces. Infra-red emission are tackled by shielding hot, visible components in order to reduce infra-red emissions generally and by means of introducing cooler air often ambient air.

Low radar signature design approaches are employed in known aircraft designs such as the Lockheed Martin F-1 17 and the Northrop Grumman B-2, although the design results look somewhat dissimilar because one aircraft has a very angular appearance in contrast to the other's smooth blended form. However, both aircraft have to some degree sacrificed engine performance for enhanced stealthiness by adopting an approach of housing the propulsion engines completely within the aircraft fuselage at the ends of inlet and outlet ducts. The configuration adopted does not easily lend itself to either thrust augmentation, by way of a reheat system, or thrust vectoring for improved manoeuvrability.

Ideally a stealthy aircraft design would combine the possibilities of either or both thrust augmentation and thrust vectoring with low radar signature and low infra-red emissions, but existing configurations as described briefly above are incompatible with such features. The present invention therefore seeks to address this objective by providing a novel nozzle shroud.

According to the broadest aspect of the invention, therefore, a nozzle shroud for an aircraft gas turbine propulsion engine exhaust nozzle comprises a structure defining duct means, open at either end into which, in use, the propulsion engine exhaust nozzle discharges hot exhaust gas, and means for inducing a stream of ambient air into the interior of the duct means to reduce the infra-red signature of the nozzle and its discharge plume.

According to another aspect of the invention the structure of the nozzle shroud is adapted to be mounted-on/dismounted-from an aircraft airframe adjacent to a propulsion engine exhaust nozzle.

The nozzle shroud will now be described in greater detail with reference, by way of example, to one embodiment thereof illustrated in the accompanying drawings, in which:

Figure 1 shows a perspective view of a nozzle shroud according to the invention, Figure 2 shows an end view of the nozzle shroud of Figure 1 in the direction of arrow A in Figures 1 and 3, Figure 3 shows a side view of the nozzle shroud in the direction of arrow B in Figure 2, Figure 4 shows a plan view of the nozzle shroud fitted in position on the rear fuselage of an aircraft.

In a conventionally laid-out airplane such as for example the Eurofighter or the Panavia Tornado, both twin-engined, the propulsion nozzles are positioned at the rear of the aircraft fuselage in ambient air. Neither aircraft incorporates radar signature reducing nor infra-red emission reducing features involving the engine exhaust nozzles. For the younger of these exemplary designs a limited amount of exhaust vectoring exists as a future possibility. The nozzle shroud of the present invention permits use of a radar and infra-red emission reducing shroud in conjunction with conventionally located exhaust nozzles, thereby permitting the possible use as desired of exhaust reheat and/or thrust vectoring.

Figure 1 shows a nozzle shroud 2 viewed from a rear quarter and above looking forwards, so that the exhaust exit apertures, upper surface and one side are visible. In Figure 4 rear portions of the aircraft fuselage 4 are outlined including the rear horizontal stabilisers 6 tails fins 8 and twin engine exhaust nozzles 10, together with the exit shroud 2 showing in a plan view their general arrangement and relative positions. Figures 2 and 3 show respectively a forward looking view of the shroud only in the direction of arrow 11 and a side view in the direction of arrow III both in Figure 4.

The nozzle shroud 2 is designed and constructed as a single-piece dismountable component normally fitted to the rearmost part 4 of an aircraft fuselage and surrounds the propulsion engine exhaust nozzle or nozzles 10.

As will be self-evident from the drawings the presently described shroud is configured for use with a twin-engined aircraft in which the two engines are mounted side-by-side in the rear of the fuselage. The illustrated shroud design consists basically of a short open tube or duct which is fitted to the rear of an aircraft fuselage surrounding the propulsion engine exhaust nozzles and extends downstream thereof by a short distance. The downstream lip 12 of the shroud 2 is angled in accordance with the stealth rules mentioned above. The upstream margin 14 of the shroud 2 is shaped according to the same rules. It is also formed with an inwardly upstanding flange 16 (Figure 1) which extends in effect around the whole of the periphery by which it may be secured to a corresponding flange (not shown) on the rear of the aircraft fuselage by means of attachment bolts or screws for example. The said fuselage mounting flange 16 extends around substantially the whole circumference of the fuselage envelope immediately adjacent the inner skin surface. The aircraft flange 18 surrounds the downstream end of the engine jet pipes, or nozzles as appropriate, and marks the end of the fuselage engine bay 20. In the case of a twin-engined aircraft the engine bay 20 is designed to accommodate the propulsion engines side-by-side and has a fire wall or dividing bulkhead 22 extending the length of the bay between the engine. This bulkhead or wall 22 terminates flush with the adjacent portions of the mounting flange 16. In this particular case the upstream margin 14 of the shroud 2 in plane view has a generally V-shaped outline and the shroud mounting flanges 16 are formed along the edges. It follows, therefore, that the corresponding fuselage mounting flange 18 must have the same, matching profile. However, for structural design reasons this particular shape may not be the best profile: for example it may be preferably to form the rearward fuselage flange in a plane perpendicular to the longitudinal axis of the aircraft, and incidentally the firewall 22. Then the mounting flanges 16,18 would not coincide with the outline of the shroud margin 14 which must retain its angled edges. Thus the mounting flange 16 on the shroud 2 may be spaced back from and angled with respect to the edge of the shroud margin 14.

The shroud 2, in this embodiment, has an upper portion 24 and a mirrorimage lower portion 26 which taken together form an outer surface which is very nearly an ellipsoid. However, its transverse section is broadened to accommodate the side-by-side layout of the engine nozzles 10. Thus, the upper and lower portions 24,26 are flattened towards the longitudinal centreline, but are. curved towards the sides. The sides may be a continuous curve but in the particular embodiment in order to accommodate the horizontal stabilisers 6 at the rear of the fuselage a rearwardly tapering slot is formed between edges 28,30 on the upper and lower shroud portions 24,26 respectively. At the downstream end of the shroud sides the upper and lower portions are joined n this particular case by angled facets 29,31, in order to closely integrate with the airframe.

Along the fuselage centreline the shroud 2 contains a longitudinal centre wall 32 extending between the upper and lower surfaces 24,26 which bisects the interior volume of the shroud in the manner as the engine bay centre fire wall 22. The downstream edges 34,36 of the upper and lower shroud portions are again angled with respect to the centre line of the fuselage and shroud and form substantially V-shaped peaks, the apexes of which lie at either end of the shroud centre wall 32. The downstream edge 33 of centre wall 32 together with upper and lower edges 34,36 and finally side edges 39 belonging to side facets 38 define substantially rectangular exit apertures in shroud 2.

In the forward looking axial view of Figure 2 the nearly rectangular exit apertures are clearly visible and the relative position of the engine propulsion nozzles 10 is also shown. It will be apparent from the indicated position of nozzles 10 in Figures 3 and 4 that these are contained completely within the envelope of shroud 2; also that the internal dimensions of the shroud leave substantial spaces surrounding the nozzles. For the dual purposes of cooling the nozzles 10 and reducing or masking infra-red emissions from the nozzles and the hot exhaust gas flow a flow of ambient air is introduced into the interior space of the shroud surrounding the nozzle etc.

A flow of tertiary air may be provided through the spaces in the rear of the fuselage surrounding the engine jet pipes to assist this cooling function. Such a flow of cooling air through the engine bay, over the exterior of the engine is well known but rarely has sufficient heat capacity to significantly affect emissions from the exhaust nozzle or the exhaust plume. In accordance with the objective of the present invention, therefore, means is provided to substantially increase a tertiary flow through the nozzle shroud. The illustrated embodiment is provided with additional apertures 40,42 in the upper and lower shroud surfaces 24,26 close to the upstream margins 16. In order to minimise emissions through these apertures and radar reflections therefrom the apertures 40,42 are preferably covered with sets of fine louvres 44,46 respectively covered with radar absorbent material. According to previous practice the edges of these apertures 40,42 and the louvres themselves are angled relative to the centreline. In the illustrated example the aperture edges and louvres form a chevron pattern and obscure the interior of the shroud and the nozzles from direct line of sight from any side location.

The louvres comprise a plurality of slots arranged parallel to the aperture edges and spaced apart each from its neighbour by a short distance. The slots are also angled relative to the shroud surface 24,26 so that air travelling rearwardly may be inducted through the spaces into the interior of the shroud with minimum loss of momentum. This is much the same as an effect found in a particular kind of nozzle known as an "ejector" nozzle. The mass flow of ambient air through these apertures is substantial.

In operation of the engines the exhaust flow through the downstream portion of the shroud will establish a pressure gradient in the interior of the shroud the effect of which is to create a continuous pumping action inducting air through the louvred apertures. This flow of ambient air through both upper and lower shroud portions surrounds the nozzles 10 with a continuously replenished cool airflow which not only cools the outer surface of the nozzle components but envelopes the hot exhaust plume in a sheath of cooler air thereby significantly reducing infra-red emissions. This tertiary flow of air is normally to be expected to be substantially in excess of any tertiary flow through the engine bay. This latter flow is also inducted into the nozzle and may benefit from pressure conditions pertaining within the shroud 2.

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Claims (4)

1 A nozzle shroud for an aircraft gas turbine propulsion engine exhaust nozzle comprises a structure defining duct means, open at either end into which, in use, the propulsion engine exhaust nozzle discharges hot exhaust gas, the nozzle being positioned relative to the nozzle shroud to obscure line of sight over most viewing angles, and means for inducing a stream of ambient air into the interior of the duct means to reduce the infra-red signature of the nozzle and its discharge plume.

2 A nozzle shroud as claimed in claim 1 wherein the structure is adapted to be mounted-on/dismounted-from an aircraft airframe adjacent to a propulsion engine exhaust nozzle.

3 A nozzle shroud as claimed in claim 1 or claim 2 adapted for use on a multiengine aircraft wherein two or more propulsion engine exhaust nozzles discharge into the interior duct.

4 A nozzle shroud as claimed in any preceding claim wherein all edges and surfaces of the shroud are arranged and adapted in accordance with principles and practices for the minimisation or elimination of radar reflection. 5 A nozzle shroud as claimed in any preceding claim wherein the means for inducing the stream of ambient air into the interior of the duct means comprises aperture means formed in the shroud walls. 6 A nozzle shroud as claimed in claim 5 wherein the or each aperture is obscured by a conductive grille. 7 A nozzle shroud as claimed in claim 5 or claim 6 wherein the conductive grille comprises a grating of conductive members.

8 8 A nozzle shroud as claimed in claim 5 or claim 6 wherein the grille comprises a mesh of conductive members. 9 A nozzle shroud as claimed in any preceding claim adapted for use on a multiengine aircraft whereby the duct means has a plurality of internal passages into which the propulsion engines individually discharge. 10 A nozzle shroud as claimed in any preceding claim wherein the structure is comprised of a lightweight metal or composite material. 11 A nozzle shroud as claimed in claim 10 wherein the lightweight material comprises a honeycombed or cellular metal. 12 A nozzle shroud substantially as hereinbefore described with reference to the accompanying drawings.