GB2038425A

GB2038425A - Gas Turbine Engine

Info

Publication number: GB2038425A
Application number: GB7849998A
Authority: GB
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1978-12-27
Filing date: 1978-12-27
Publication date: 1980-07-23
Also published as: GB2038425B

Abstract

The engine has an axial flow compressor or fan provided with a second compressor built into the nose bullet 20 of the engine for energizing the air 42 which flows over the roots of the blades 40 of the first compressor. The second compressor comprises an annulus 36 defined by a cowl 24 which surrounds and is attached to the nose bullet 20 by hollow vanes 34 through which anti- icing air flows. <IMAGE>

Description

SPECIFICATION Gas Turbine Engine This invention concerns gas turbine propulsion engines having axial flow compressors. More particularly, but not exclusively the invention concerns gas turbine propulsion engines of the kind which include a ducted fan, from which the majority of the propulsive thrust is obtained.

All axial flow compressors have certain common characteristics i.e. they comprise a number of stages of rotor blades mounted on respective discs, each rotor stage being adjacent to and upstream of, a stator stage. Each rotor blade lies on its own radial line, radial that is with respect to the axis of rotation of the compressor as a whole. The outer extremity of each blade rotates as a faster peripheral speed than its hub or root and moreover, the outer extremity of each blade is twisted more than its root, relative to the direction of rotation. It follows that a rotor blade for an axial flow compressor cannot impart as much energy to air flowing over its root, as it can to air flowing over the remainder of its length.

Efforts have been made to energise the air flow over the root portions of the rotor stages in an axial flow compressor and, more recently, similar efforts have been made, to solve the same problem which also exists in the fan of a ducted fan, gas turbine engine. A construction which achieves the desired object is described and claimed in our published patent specification 1,257,497. However, one embodiment of that construction requires stator stages which rotate as well as rotor stages, in order to drive the main fan and all embodiments therein require an additional fan stage to provide the energy in the air at the fan blades roots.

The present invention seeks to provide a construction which in operation of a gas turbine engine to which it is fitted, will energise air prior to directing it over the roots of an adjacent stage of fan compressor rotor blades, which construction does not comprise a disc and rotor blade constructions, but is easier and cheaper to manufacture relative to a disc and rotor blade construction.

According to the present invention there is provided a gas turbine engine including first and second compressors wherein the first compressor comprises a rotatable nose bullet surrounded by a cowl to form a compressor annulus therebetween and separated from the cowl by a plurality of equi-angularly spaced vanes which define diffusing passages for ambient airflow, the outlet of said first compressor being aligned with, the roots of the first stage of rotor blades of the second compressor so as to pre-energise said air flow prior to ejecting it over said roots.

The second compressor may comprise the compressor of a gas turbine jet propulsion engine.

Alternatively the second compressor may comprise the fan of a ducted fan, gas turbine engine.

Preferably the nose bullet is rotated by a turbine of the engine.

The nose bullet may be connected for driving by said turbine via a rotor stage of said second compressor of fan. Alternatively, the nose bullet may be connected for driving by said turbine via a stepping gear box.

The invention will now be described, by way of example, with reference to the accompanying drawings in which: Figure 1 is a diagrammatic view of a gas turbine engine incorporating a nose bullet in accordance with an embodiment of the invention.

Figure 2 is an enlarged, cross sectional part view of Figure 1, Figure 3 is an enlarged, cross sectional part view of Figure 1, including a further embodiment of the invention.

In Figure 1 a gas turbine engine is generally indicated by the numeral 10. Gas turbine engine 10 comprises an axial flow compressor 12, combustion equipment 14, a turbine 1 6 by which compressor 12 is driven and an exhaust nozzle 18.

A nozzle bullet 20 is mounted at the extreme upstream end of engine 10, and is connected for coaxial rotation with compressor 12 as is described hereinafter. Nose bullet 20 comprises a central portion 22 and a cowl 24 which is spaced from the central portion to form an annulus therebetween.

Referring now to Figure 2 in the present example, the central portion 22 of nose bullet 20 comprises a double walled structure located at its middle on a shaft 26, the other end of which (not shown) is drivingly connected to turbine 1 6 (not shown in Figure 2). A cap nut 28 retains portion 22 on shaft 26.

The downstream extremity of central portion 22 terminates in an annular flange 30 via which the nose bullet 20 is fastened by setscrews 31 to the rotor disc 32 of the first rotor stage of compressor 1 2 for co-rotation therewith.

Cowl 24 is spaced from central portion 22 by a plurality of equiangularly spaced vanes 34. Vanes 34 are arranged in the manner of rotor blades in a single stage, axial flow compressor i.e. on rotation of the nose bullet, air is induced into the annulus 36 between cowl 24 and central portion 22 in the direction of arrows 38 and, on entering the passages formed by pairs of adjacent vanes, is accelerated and diffused. Energy is thus imparted to the airflow through annulus 36. The annulus profile is a constricting one in a downstream direction as in normal compressor design and this - has appropriate affect on the energy imparted to the airflow through it.

The energised air is ejected from annulus 36 directly onto the root portions of the first stage rotor blades 40, as indicated by arrows 42, thus to some extent, improving the efficiency of the rotor blades over their root portions and ensuring a more uniform air pressure and velocity distribution over the full length of each rotor blade 40.

The double wall construction of central portion 22 of nose bullet 20 enables anti-icing air to be directed to the interior of cowl 23 via the interior of turbine shafts 26 and the space between the double walls 22a, 22b of portion 22. The antiicing air then passes radially outwards through hollow vanes 34 into the interior of cowl 24 and exits from slots 44 in the inner lip 46 of the cowl.

The pre-heated anti-icing air, when ejected into the annulus 36, heats the air taken from ambient atmosphere by the rotary actions of nose bullets 20, thus supplementing the heat generated by vanes 34 doing work on the air flowing between them.

In some engine designs, it is necessary to provide a stage of fixed, inlet guide vanes upstream of the first rotor stage of a second compressor as defined herein. Such an arrangement is shown in Figure 3, where inlet guide vanes 44 are immediately upstream of compressor rotor stage 46.

The radially inner end 48 off each inlet guide vane 44 is provided with a pair of platforms 50, 52 so arranged with respect to each other, that the complete stage of guide vanes 44 forms a diffusing annulus 54.

The diffusing annulus 54 forms a continuation of diffusing annulus 56 which in turn is defined by cowl 58 and central portion 60, both of which are rotatahle in the manner of their counterparts in Figure 2. Air energised by diffusing annulus 56 is at least maintained in its energised state by diffusing annulus 54 prior to being ejected over the roots 62 of the compressor blades of rotor stage 46.

The leading edges 64 of struts 66 which separate cowl 58 and central portion 60, may be curved in the direction of rotation, to improve air intake efficiency. Similarly, the trailing edges 66 of struts 68 which separate inner and outer platforms 52, 50 may be curved, to eject energised air in an appropriate direction for entry into the rotor stage 46.

Claims

1. In a gas turbine engine having a first axial flow compressor comprising one or more rotors having radially extending angularly spaced compressor blades, the provision of a second compressor located upstream of the first compressor, said second compressor comprising a rotatable nose bullet surrounded by a cowl to form a compressor annulus therebetween and separated from the cowl by a plurality of equiangularly spaced vanes which define flow passages for ambient airflow, the outlet of said second compressor being aligned with, the roots of the compressor blades of the first stage of the first compressor so as to pre-energise said air flow prior to ejecting it over the roots.

2. A gas turbine engine according to claim 1 wherein the first compressor comprises a fan mounted for rotation in a by-pass duct of the engine.

3. A gas turbine engine according to claim 1 wherein the nose bullet is connected to a turbine of the engine to be driven thereby.

4. A gas turbine engine according to claim 1 wherein the flow passages are diffusing passages.

5. A gas turbine engine according to claim 3 wherein the nose bullet is connected to a rotor of a compressor which is driven by the turbine.

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6. A gas turbine engine according to claim 5 wherein the second compressor is connected to, and rotates with, the first compressor.

7. A gas turbine engine according to claim 3 wherein the nose bullet is connected to the turbine by means of a gear box.

8. A gas turbine engine substantially as hereindescribed with reference to the drawings filed with application No. 49998/78.