COOLED PUSHER PROPELLER SYSTEM

This application is a divisional of U.S. patent application Ser. No. 12/030,400, filed Feb. 13, 2008.

The present invention relates to a propeller system, and more particularly to a propeller system for a pusher type counter-rotating propulsion system.

Turboprop engines are commonly designed to drive either a single row of propellers or two rows of counter rotating propellers. The propeller(s) may be mounted forward of the engine (“tractor” installation) or rearwardly of the engine (“pusher” installation). In the pusher arrangement, the engine has an efficient free stream inlet, and the high speed propeller jet does not impinge on airplane or nacelle surfaces, thus avoiding scrubbing drag. Also, a pusher engine installation on the aircraft empennage minimizes cabin noise generated by wing-mounted tractor nacelle engines.

A pusher arrangement, however, may complicate the location for the gas turbine exhaust. One pusher arrangement locates the gas turbine exhaust upstream of the propellers in an annular or lobed configuration. This configuration may require the nacelle skin downstream of the exhaust and a root section of the propeller blades to be designed for elevated temperatures caused by the hot exhaust gases. Furthermore, hot engine exhaust directed past the pusher propeller may increase noise and reduce airfoil life.

A propulsion system according to an exemplary aspect of the present invention includes: a first row of propeller blades which rotate about an axis; an annular cooling flow nozzle about an axis, the annular cooling flow nozzle upstream of the first row of propeller blades; and an annular exhaust nozzle about an axis radially outboard of the annular cooling flow nozzle, the annular exhaust nozzle upstream of the first row of propeller blades.

A method of directing an exhaust flow from a propulsion system according to an exemplary aspect of the present invention includes: ejecting a cooling flow through an annular cooling flow nozzle upstream of a first row of propeller blades which rotate about an axis; and ejecting an exhaust flow through an annular exhaust nozzle about an axis radially outboard of the annular cooling flow nozzle.

FIG. 1A schematically illustrates a pusher type counter-rotating propulsion system 10. A gas turbine engine 12 generally includes a compressor section A, a combustor section B and a turbine section N as generally understood. Air is compressed by the compressor section A, mixed and burned with fuel within the combustor section B, then expanded over the turbine section T to generate a high temperature exhaust gas flow E. The exhaust gas flow E from the turbine section N of the gas turbine engine 12 is communicated through an annular exhaust nozzle 14 (FIG. 1B) upstream of a first row of propeller blades 16A which rotate about an axis X. A second row of row of propeller blades 16B may be located downstream of the first row of propeller blades 16A in a counter-rotating propeller (CRP) propfan pusher configuration about the axis X.

Referring to FIG. 1B, the annular exhaust nozzle 14 in one non-limiting embodiment is defined radially outboard of an annular cooling flow nozzle 18. A cooling flow C and/or other airflow that is different from the exhaust gas flow E is communicated through the annular cooling flow nozzle 18. The cooling flow C may be sourced from, for example only, a compressor section of the gas turbine engine 12, an inlet or other source.

The annular exhaust nozzle 14 and annular cooling flow nozzle 18 form an annular nozzle section 15 generally upstream of the first row of propeller blades 16A. The annular exhaust nozzle 14 and annular cooling flow nozzle 18 may be radially defined between a radially outboard engine nacelle 23 and a radially inboard first spinner section 20A which mounts the first row of propeller blades 16A (FIG. 1B). A second spinner section 20B which mounts the second row of propeller blades 16B is downstream of the first spinner section 20A. The second spinner section 20B counter rotates relative the first spinner section 20A while a tail cone 21 is located downstream of the second spinner section 20B. The tail cone 21 is aerodynamically shaped and in one embodiment remains rotationally stationary to facilitate flow.

The nozzles 14, 18 have a relatively low profile and are aerodynamically shaped to minimize noise, drag and weight. The exhaust gas flow and the cooling flow are substantially uniform and directed through the first and second row of propeller blades 16A, 16B to avoid significant noise addition. The nozzles 14, 18 direct the exhaust stream E and the cooling flow C aftward such that at least a portion of the propulsive energy within the streams are exploited.

Referring to FIG. 2, the cooling flow C is directed essentially under the annular exhaust gas flow E to enter each blade 22 at a root section 24 thereof. It should be understood that although only a single propeller blade 22 will be described, the description is similarly applicable to each of the propeller blades 22 in the first row of propeller blades 16A and/or the second row of propeller blades 16B. The cooling flow C operates to lower the temperature of the root section 24 which is proximate the exhaust gas flow E as well as film cool and insulate the spinner sections 20A, 20B and tail cone 21 at least partially from the exhaust gas flow E. The root section 24 may additionally be mechanically insulated or otherwise hardened to further resist the exhaust gas flow E.

The root section 24 includes an intake 26 which communicates with a distribution channel 28. The distribution channel 28 may be defined at least in part by hollow airfoil sections and/or passages in the blade 22. The cooling flow C is communicated through the distribution channel 28 to an exit 30. Once the cooling flow enters the blade 22, the cooling flow C is self-pumped, by the rotational force of blade 22, through the distribution channel 28 which results in thrust recovery when ejected from the exit 30. The exit 30 may be located adjacent a trailing edge section 32 and/or a blade tip section 34. The exit 30 may alternatively or additionally be distributed along the blade trailing edge section 32 to minimize an airfoil wake deficit with a corresponding potential noise reduction. The cooling flow may additionally be directed through a forward distribution channel section 28A adjacent a leading edge section 36 where hot gas impingement exists so as to reduce noise generation while the annular exhaust flow path is generally maintained.

It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.

It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit from the instant invention.

Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention.

The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The disclosed embodiments of this invention have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.