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EP-3770415-B1 - GEARED TURBOFAN ENGINE WITH HIGH COMPRESSOR EXIT TEMPERATURE

EP3770415B1EP 3770415 B1EP3770415 B1EP 3770415B1EP-3770415-B1

Inventors

  • SCHWARZ, FREDERICK M.
  • HASEL, KARL L.

Dates

Publication Date
20260506
Application Date
20130312

Claims (7)

  1. A gas turbine engine (20) comprising: a fan (42) including a plurality of fan blades rotatable about an axis; a compressor section (24) including at least a first compressor section (44) and a second compressor section (52); a combustor (26) in fluid communication with the compressor section (24); a turbine section (28) in fluid communication with the combustor (26); and a geared architecture (48) driven by the turbine section (28) for rotating the fan (42) about the axis, wherein: the compressor section (24) is configured to operate at an average exit temperature between about 1000 °F (538 °C) and about 1500 °F (816 °C), wherein the operating temperature is defined at Sea Level, end of takeoff power and at a rated thrust for the gas turbine engine; the first compressor section comprises a low pressure compressor (44) and the second compressor section comprises a high pressure compressor (52), and wherein the turbine section (28) comprises a low pressure turbine (46) that drives the low pressure compressor (44) via a first shaft (30) and a high pressure turbine (54) that drives the high pressure compressor (52) via a second shaft (32), and wherein the geared architecture (48) couples the first shaft (30) to the fan (42); and the high pressure compressor (52) includes a plurality of stages with each stage comprising a disk (94) with a plurality of blades (96) extending radially outwardly from a rim of the disk (94), and wherein the plurality of stages includes at least a first stage located at the front of the high pressure compressor (52) having a first blade and disk configuration and a second stage located at the rear of the high pressure compressor (52) having a second blade and disk configuration that is different than the first blade and disk configuration; characterised in that the first blade and disk configuration comprises a plurality of slots to receive the plurality of blades and including a plurality of rim cavities (98) for honeycomb seals, and wherein the second blade and disk configuration comprises integrally formed blades such that there are no rim cavities or associated honeycomb seals.
  2. The gas turbine engine according to claim 1 wherein the average exit temperature is between about 1100 °F (593 °C) and about 1450 °F (788 °C).
  3. The gas turbine engine according to claim 1, wherein the average exit temperature is between about 1450 °F (788 °C) and about 1500 °F (816 °C).
  4. The gas turbine engine according to any preceding claim, wherein the fan (42) drives air along a bypass flow path (B) in a bypass duct defined between a fan nacelle (15) and a core nacelle (16), and wherein a bypass ratio is greater than about ten.
  5. The gas turbine engine (20) according to any preceding claim wherein the geared architecture (48) has a gear ratio that is greater than about 2.4.
  6. The gas turbine engine according to any preceding claim, wherein: the first compressor section (44) rotates at a first speed and the second compressor section (52) rotates at a second speed greater than the first speed, and wherein the average exit temperature of the second compressor section (52) is between about 1000 °F (538 °C) and about 1500 °F (816 °C).
  7. The gas turbine engine according to any preceding claim, wherein the low pressure compressor (44) includes bladed disks.

Description

BACKGROUND A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds. Under certain operational conditions, the compressor section of geared turbofan engines can be subjected to high exit temperatures. Although geared turbofan architectures have provided improved propulsive efficiency at high temperatures, turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies. Bangalore Philip R. Gliebe ET AL, "Ultra-High Bypass Engine Aeroacoustic Study" NASA/CR-2003-212525, 1 October 2003, XP 055277347 discloses a prior art gas turbine engine as set forth in the preamble of claim 1. US 2009/053058 A1 discloses a prior art gas turbine engine with axial movable fan variable area nozzle. SUMMARY From a first aspect, there is provided a gas turbine engine as recited in claim 1. Features of embodiments of the invention are set forth in the dependent claims. BRIEF DESCRIPTION OF THE DRAWINGS The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: Figure 1 schematically illustrates a geared turbofan engine embodiment.Figure 2 illustrates a compressor section of the engine of Figure 1.Figure 3 is a chart showing operating conditions for three different engines having high compressor exit temperatures that are operable with the subject invention.Figure 4 illustrates a blisk with a rim cavity configuration.Figure 5 illustrates a blisk without rim cavities.Figure 6 is a schematic representation of a compressor section with an inducer. DESCRIPTION Figure 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B in a bypass duct defined between a fan nacelle 15 and a core nacelle 16, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application. The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mi