Integrated fan / low pressure compressor rotor for gas turbine engine

• US 6,145,300 A
• Filed: 07/09/1998
• Issued: 11/14/2000
• Est. Priority Date: 07/09/1998
• Status: Expired due to Term

• Current Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Romani, Giuseppe
• Primary Examiner(s): Thorpe, Timothy S.
• Assistant Examiner(s): TORRENTE, DAVID JOHN
• Application Number:
  US09/112,236
• Time in Patent Office: 859 Days
• Field of Search: 60/226.1, 60/226.2, 60/226.3, 60/262, 415/199.4, 416/201 R, 416/198 A, 416/200 A, 416/203
• US Class Current: 60/226.1
• CPC Class Codes: F01D 5/141 Shape, i.e. outer, aerodynamic form F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147
• CPC Class Codes: F01D 5/142 of the blades of successive rotor or stator blade-rows
• CPC Class Codes: F02K 3/06 with front fan
• CPC Class Codes: F04D 29/324 Blades
• CPC Class Codes: F04D 29/325 for axial flow fans blade mountings F04D29/34, blades F04D29/38
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

An integrated fan and low pressure compressor rotor having a hub with a circumferential array of primary blades. Each primary blade has a forward fan blade portion housed in the engine inlet radially extending from the hub to the fan case and a rearward compressor blade portion housed in the core duct radially extending from the hub to the intermediate case. The fan blade portion and compressor blade portion have aerodynamically aligned lateral airfoil surfaces and platforms such that the fan blade portion and compressor blade portion have continuously merged lateral airfoil surfaces and platforms. The integrated rotor replaces two conventional rotors resulting in cost savings.

Claims

1. In a turbofan gas turbine engine having:

a fan case;

an intermediate case with a forward flow splitter; and

an engine core with a shaft mounted for rotation about a longitudinal axis, an annular bypass duct defined between the fan case and intermediate case, an annular core duct defined between the core and the intermediate case, and an engine inlet defined inward of the fan case forward of the flow splitter, the improvement comprising an integrated fan and low pressure compressor rotor comprising;

a hub mounted to a forward end of the shaft, the hub having a circumferential array of primary blades, each primary blade having;

a forward fan blade portion housed in the engine inlet radially extending from the hub to the fan case; and

a rearward compressor blade portion housed in the core duct radially extending from the hub to the intermediate case, the fan blade portion and compressor blade portion having aerodynamically aligned lateral airfoil surfaces and platforms wherein the fan blade portion and compressor blade portion have continuously merged lateral airfoil surfaces and platforms.

2. An integrated fan and low pressure compressor rotor according to claim 1 wherein the fan blade portion and compressor blade portion are releasably fastened together.

3. An integrated fan and low pressure compressor rotor according to claim 1 further comprising a circumferential array of auxiliary compressor blades disposed between compressor blade portions of the circumferential array of primary blades and housed in the core duct radially extending from the hub to the intermediate case.

4. An intake for a turbofan gas turbine engine comprising:

an engine support structure comprising;

a fan case with external engine mounts disposed in a radially extending support surface defined by the mounts and a point on a longitudinal axis of the engine;

an intermediate case with a forward flow splitter;

engine core support means for mounting a shaft for rotation about the longitudinal axis;

an annular bypass duct defined between the fan case and intermediate case;

an annular axially cascading array of bypass stator blades aligned in the support surface and spanning between the fan case and the intermediate case;

an annular core duct defined between an engine core and the intermediate case;

an annular axially cascading array of core stator blades aligned in the support surface and spanning between the intermediate case and the core;

an engine inlet defined inward of the fan case forward of the flow splitter;

the engine further comprising,

an integrated fan and low pressure compressor rotor immediately upstream of the engine support structure, the rotor comprising;

a hub mounted to a forward end of the shaft, the hub having a circumferential array of primary blades, each primary blade having;

a forward fan blade portion housed in the engine inlet radially extending from the hub to the fan case; and

a rearward compressor blade portion housed in the core duct radially extending from the hub to the intermediate case, the fan blade portion and compressor blade portion having aerodynamically aligned lateral airfoil surfaces and platforms wherein the fan blade portion and compressor blade portion have continuously merged lateral airfoil surfaces and platforms.

5. An intake for a turbofan gas turbine engine according to claim 4 wherein the fan blade portion and compressor blade portion are releasably fastened together.

6. An intake for a turbofan gas turbine engine according to claim 4 further comprising a circumferential array of auxiliary compressor blades disposed between compressor blade portions of the circumferential array of primary blades and housed in the core duct radially extending from the hub to the intermediate case.