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Title: Gas turbine engine with supersonic compressor

Abstract

A gas turbine engine having a compressor section using blades on a rotor to deliver a gas at supersonic conditions to a stator. The stator includes one or more of aerodynamic ducts that have converging and diverging portions for deceleration of the gas to subsonic conditions and to deliver a high pressure gas to combustors. The aerodynamic ducts include structures for changing the effective contraction ratio to enable starting even when designed for high pressure ratios, and structures for boundary layer control. In an embodiment, aerodynamic ducts are provided having an aspect ratio of two to one (2:1) or more, when viewed in cross-section orthogonal to flow direction at an entrance to the aerodynamic duct.

Inventors:
;
Publication Date:
Research Org.:
Dresser-Rand Company, Olean, NY (United States)
Sponsoring Org.:
USDOE
OSTI Identifier:
1223700
Patent Number(s):
9,163,521
Application Number:
13/542,669
Assignee:
Dresser-Rand Company (Olean, NY) NETL
DOE Contract Number:
FE0000493
Resource Type:
Patent
Resource Relation:
Patent File Date: 2012 Jul 06
Country of Publication:
United States
Language:
English
Subject:
33 ADVANCED PROPULSION SYSTEMS

Citation Formats

Roberts, II, William Byron, and Lawlor, Shawn P. Gas turbine engine with supersonic compressor. United States: N. p., 2015. Web.
Roberts, II, William Byron, & Lawlor, Shawn P. Gas turbine engine with supersonic compressor. United States.
Roberts, II, William Byron, and Lawlor, Shawn P. Tue . "Gas turbine engine with supersonic compressor". United States. doi:. https://www.osti.gov/servlets/purl/1223700.
@article{osti_1223700,
title = {Gas turbine engine with supersonic compressor},
author = {Roberts, II, William Byron and Lawlor, Shawn P.},
abstractNote = {A gas turbine engine having a compressor section using blades on a rotor to deliver a gas at supersonic conditions to a stator. The stator includes one or more of aerodynamic ducts that have converging and diverging portions for deceleration of the gas to subsonic conditions and to deliver a high pressure gas to combustors. The aerodynamic ducts include structures for changing the effective contraction ratio to enable starting even when designed for high pressure ratios, and structures for boundary layer control. In an embodiment, aerodynamic ducts are provided having an aspect ratio of two to one (2:1) or more, when viewed in cross-section orthogonal to flow direction at an entrance to the aerodynamic duct.},
doi = {},
journal = {},
number = ,
volume = ,
place = {United States},
year = {Tue Oct 20 00:00:00 EDT 2015},
month = {Tue Oct 20 00:00:00 EDT 2015}
}

Patent:

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  • A midframe portion (313) of a gas turbine engine (310) is presented and includes a compressor section with a last stage blade to orient an air flow (311) at a first angle (372). The midframe portion (313) further includes a turbine section with a first stage blade to receive the air flow (311) oriented at a second angle (374). The midframe portion (313) further includes a manifold (314) to directly couple the air flow (311) from the compressor section to a combustor head (318) upstream of the turbine section. The combustor head (318) introduces an offset angle in the airmore » flow (311) from the first angle (372) to the second angle (374) to discharge the air flow (311) from the combustor head (318) at the second angle (374). While introducing the offset angle, the combustor head (318) at least maintains or augments the first angle (372).« less
  • A midframe portion (213) of a gas turbine engine (210) is presented, and includes a compressor section (212) configured to discharge an air flow (211) directed in a radial direction from an outlet of the compressor section (212). Additionally, the midframe portion (213) includes a manifold (214) to directly couple the air flow (211) from the compressor section (212) outlet to an inlet of a respective combustor head (218) of the midframe portion (213).
  • A cooling system for a turbine engine for directing cooling fluids from a compressor to a turbine blade cooling fluid supply and from an ambient air source to the turbine blade cooling fluid supply to supply cooling fluids to one or more airfoils of a rotor assembly is disclosed. The cooling system may include a compressor bleed conduit extending from a compressor to the turbine blade cooling fluid supply that provides cooling fluid to at least one turbine blade. The compressor bleed conduit may include an upstream section and a downstream section whereby the upstream section exhausts compressed bleed airmore » through an outlet into the downstream section through which ambient air passes. The outlet of the upstream section may be generally aligned with a flow of ambient air flowing in the downstream section. As such, the compressed air increases the flow of ambient air to the turbine blade cooling fluid supply.« less
  • A gas turbine engine has an inlet compressor with a stepped centrifugal impeller and a double diffuser therefrom controlled by valve means and operative to supply air to combustion apparatus from when motive fluid is directed through a turbine to drive the compressor. The turbine may drive the load or a second power turbine in series with the compressor drive turbine can be included. The discharge from the double diffuser is controlled to vary the mass flow of the engine for operation at idling and under light loads and operable to be completely open for higher power output of themore » engine.« less
  • This patent describes a single shaft gas turbine engine power plant comprising: a power head and a radial load compressor driven by the power head, the power head being an air breathing gas turbine engine comprising a radial compressor; a combustor for heating, and thereby adding energy to, the air discharged from the radial compressor; a radial gas turbine driven by hot gases flowing thereto from the combustor, the radial gas turbine being drive connected to the engine and load compressor; and the power plant further comprising means for so scheduling a flow of fuel to the turbine engine combustormore » as to be simultaneously responsive to a radial gas turbine exhaust temperature and to gas turbine engine speed. This is needed to maintain the turbine exhaust temperature below selected maximums during start-up and acceleration of the gas turbine engine to a specified speed and to thereafter keep the gas turbine engine speed from exceeding a selected maximum.« less