Gas turbine engine having bleed apparatus with dynamic pressure recovery

• US 4,236,869 A
• Filed: 12/27/1977
• Issued: 12/02/1980
• Est. Priority Date: 12/27/1977
• Status: Expired due to Term

• Current Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Laurello, Vincent P.
• Primary Examiner(s): Powell, Jr., Everette A.
• Assistant Examiner(s): Trausch, III, A. N.
• Application Number:
  US05/864,192
• Time in Patent Office: 1,071 Days
• Field of Search: 60/39.66, 415/115, 415/116, 415/175, 416/90 R, 416/92, 416/96 R, 416/97 R, 416/95
• US Class Current: 416/95
• CPC Class Codes: F01D 5/081 Cooling fluid being directed on the side of the rotor disc or at the roots of the blades F01D5/087 takes precedence
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A gas turbine engine has conduits leading from the compressor to the turbine for bleeding a portion of compressor air through the turbine blades for cooling or other purposes. A radial pump in the conduits formed by a tangential accelerator and a rotating diffuser transmits the air from the non-rotating environment in the engine to the rotating environment of the turbine rotor. The tangential accelerator accelerates the bleed air to a tangential speed higher than that of the rotor, and the radial diffuser is designed to recover the dynamic pressure or relative velocity of the air over the rotor.

Claims

1. In a gas turine of the type having a compressor section having a flow path from which a fluid working medium such as air is discharged at an elevated pressure to a combustion section where the working medium is combined with a fuel in a combustion process and the products of combustion are directed at high velocity onto a plurality of turbine blades distributed about the periphery of a turbine rotor in a turbine section, improved apparatus for ducting the working medium to the plurality of turbine blades for cooling and other purposes comprising:

first conduit means receiving a portion of the working medium from the flow path in the compressor section and having a discharge end located adjacent the turbine rotor and forming a tangential accelerator discharging the working medium substantially tangentially of an annular region of the rotor in the direction of rotor rotation and at a higher velocity than the velocity of the rotor in the annular region at operating speed, whereby a relative velocity exists between the discharged working medium and the rotor, and the temperature of the working medium is reduced by the higher velocity for greater cooling effect, the accelerator including a plurality of tangentially directed outlets arranged in an annular array adjacent said annular region of the rotor; and

second conduit means connected with the turbine rotor for rotation with the turbine blades and leading to the turbine blades at the periphery of the rotor, the second conduit means having inlets directed generally tangentially of the rotor and positioned radially outward from the outlets of the tangential accelerator and said annular region of the rotor at the same axial station as the accelerator outlets to receive the working medium discharged at high velocity from the tangential accelerator, the tangentially directed inlets connecting to a downstream portion of the second conduit means extending radially outward on the rotor from the annular region and having diverging walls forming diffusing means for diffusing the working medium received at the higher velocity and recovering the dynamic pressure of the working medium whereby static pressure is improved in the second conduit means.

2. In a gas turbine engine, the improved apparatus of claim 1 wherein the second conduit means is comprised by a plurality of ribs on the rotor forming an annular series of the tangential inlets adjacent said annular region of the rotor and leading from the inlets radially outward of the rotor.

3. In a gas turbine engine, the improved apparatus of claim 2 wherein the ribs of the second conduit means are sandwiched between the turbine rotor and a side plate attached to the rotor.

4. The apparatus of claim 3 wherein the side plate attached to the rotor is a segmented side plate and the ribs are integrally connected to the segments of the plate.

5. The improved apparatus of claim 1 wherein the tangential accelerator is comprised of a plurality of ribs forming the plurality of outlets and having portions extending generally in the tangential direction at the discharge openings of the outlets adjacent the rotor, and radially inner and outer plates sandwiching the ribs therebetween.

6. The improved apparatus of claim 5 wherein the ribs of the tangential accelerator also include portions upstream of the discharge openings extending generally axially of the rotor and joining the tangential portions near the discharge openings.

7. The improved apparatus of claim 6 wherein the second conduit means is comprised by a plurality of flow turning ribs disposed at one side of the rotor and extending at one end tangentially of the annular region of the rotor to form the tangential inlets and bending from the tangential to the radial direction for moving the medium to the blades at the periphery of the rotor.