News Column

Patent Issued for Gas Turbine Engine with Dual Compression Rotor

June 3, 2014



By a News Reporter-Staff News Editor at Journal of Technology -- The United States of America as represented by the Secretary of the Air Force (Washington, DC) has been issued patent number 8726635, according to news reporting originating out of Alexandria, Virginia, by VerticalNews editors.

The patent's inventor is Dale, Mark R. (Beavercreek, OH).

This patent was filed on December 17, 2012 and was published online on May 20, 2014.

From the background information supplied by the inventors, news correspondents obtained the following quote: "Gas turbine engines are well-known for providing propulsion and power production. Compared to other engine types, a gas turbine engine has the advantage that its rotors undergo purely rotational motion, and it can therefore operate at high speed with minimum vibration. In such engines, the theory of operation is described thermodynamically by the Brayton cycle: air is compressed isentropically, combustion of air/fuel mixture occurs, and expansion over turbine blades occurs isentropically back to the starting pressure. Such engines, however, typically operate efficiently only within a relatively narrow band of engine speeds.

"Conventional modern turbine engines have to run at high temperatures to produce enough work to drive multiple stages of compression to achieve high pressure ratios and high efficiency. The high temperatures and pressures drive up design complexity and life cycle costs (development, production, maintenance). Also, these turbine engines are designed for optimum performance at one design point which causes the engines to operate less efficiently during off design conditions. That is, conventional turbine engines cannot produce the optimum higher pressure ratios at lower cruise power settings.

"Another limitation of conventional turbine engines is the use of a rotor for each compressor and turbine stage, as well as a connecting shaft linking the compressor and turbine, which significantly increases the engine length, weight, and cost. Also, a limiting factor is the ability of steel, nickel, ceramic, or other engine materials to withstand extreme heat and pressure. These extreme temperatures require elaborate/complex secondary flow cooling circuits to maintain acceptable material properties, especially at the high pressure turbine bores. Unfortunately, these cooling systems reduce engine performance and add undesirable weight. Even when cooling systems are used in turbine engines, there is centrifugal and thermal growth of the rotors. Maintaining a gap between the rotor tips and the engine shroud to account for thermal growth causes large tip leakage losses and lower component performance and efficiency. Conventional turbine engines also have complex bearing systems that operate near maximum temperature limits and endure extreme shaft dynamics.

"There are a myriad of known gas turbine engine configurations. One early example is illustrated in FIG. 1. The engine in FIG. 1 was designed by Hans von Ohain in 1937 and was designated the He.S3 turbojet engine. The combustor of the Ohain engine is positioned in the large unused space in front of the radial-flow compressor. Airflow through the He.S3 engine followed a generally S-shaped configuration.

"U.S. Pat. No. 2,694,291 to Rosengart describes a rotor and combustion chamber arrangement for gas turbines. The turbine includes a stationary combustion chamber that is generally toroidal shaped and has a continuous opening at its inner periphery. Mounted on the rotor are hollow blades designed so that air and exhaust gases pass between the blades while cooling air passes within the hollows of the blades to facilitate cooling.

"Another gas turbine example is described in U.S. Pat. No. 3,269,120 to Sabatiuk. Sabatiuk discloses a gas turbine engine having compressor and turbine passages in a single rotor element. The engine has axial flow compressor passages and radial flow turbine passages in a single rotor. Flow through the compressor passages is in a direction generally parallel to the axis of the rotor element, and flow through the turbine passages is in a radial direction at least for a portion of the length of the passages.

"U.S. Pat. No. 3,892,069 to Hansford describes a propulsion unit for an aircraft that includes rotor means incorporating a multi-bladed fan which over an outer peripheral region thereof defines centrifugal flow compressor passages and centripetal flow turbine passages and an annular combustor encircling the rotor means. The combustor has inlet means for directing air from the compressor passages into the combustor and outlet means for directing combustion gases from the combustor into the turbine passages to drive the rotor means. Hansford's propulsion unit includes an air intake leading to a series of circularly distributed centrifugal flow compressor passages and includes an annular combustor which is of substantially toroidal shape and defines a combustion chamber of part circular cross-section.

"In addition, U.S. Pat. No. 6,988,357 to Dev discloses a gas turbine engine including a combustion chamber section, a turbine section, and a compressor section. The turbine section surrounds the combustion chamber, and the compressor section surrounds the turbine section. The rear rotor of the turbine engine includes an integral compressor section on the outside and a turbine section on the inside.

"Given the limitations of conventional gas turbine engines, there is a need for an improved engine that minimizes weight and fuel consumption while maximizing thrust and efficiency."

Supplementing the background information on this patent, VerticalNews reporters also obtained the inventor's summary information for this patent: "The gas turbine engine of the present invention (also referred to as the Revolutionary Innovative Turbine Engine (RITE)) incorporates technology that provides versatile mission capabilities in one configuration, including: maximum engine thrust for minimum takeoff and climb; minimum engine fuel consumption and vehicle drag for maximum range; and maximum power generation to meet increasing vehicle and weapon requirements.

"The RITE cycle has infinite variability and applicability because each rotating stage can be controlled independently. The technology features that make up the RITE cycle include, but are not limited to: a dual compression rotor (DCR) for higher overall pressure ratios with minimum number of rotor stages; independently supported rotors (ISR) using hybrid and/or magnetic bearings for support in lieu of shafts; axial and centrifugal compressors controlled independently for optimum performance match; substantial reduction in overall engine length and weight; a dedicated turbine attached to each fan/compressor rotor (cooling air on-board if needed to control turbine metal temperatures (cooling air provided directly to turbine through radial passages in DCR); low turbine bore temperatures due to cool compressor inlet airflow bathing; hot sections located outboard for overall reduction of internal temperatures for bearings and mechanical systems; reduced thermal and centripetal bore growth for tighter turbine tip clearance; use of nickel material at high compressor exit temperature region; and inter turbine burners (ITB) to boost temperature and work capability after each turbine to lower maximum turbine temperature and net energy loss throughout the engine.

"Other features include: fan-on-turbine which eliminates the need for a low pressure shaft; an internal starter for engine starting; generators that can be controlled by controlling the speed of each rotor separately for optimized operability (stage matching, power extraction and management (all electric, no gearbox required)); engine core can run at design point throughout mission profile because thrust requirements for flight can be independently controlled by the ITBs, greatly reducing engine thermal cycles and maintenance costs and increasing engine life; and a free/power turbine that matches inlet airflow to reduce vehicle drag (at lower power settings during vehicle cruise or loiter (the power turbine temperature can be independently increased to match inlet design flow by compensating for decreased fan flow). The power turbine drives a power generator which maximizes the use of gas path energy/efficiency, extracts energy to reduce engine gas temperature, and reduces power generation impact on the engine cycle. At lower power settings during vehicle cruise or loiter the power turbine temperature can be independently increased to maximize power extraction for avionics or directed energy weapons.

"In one exemplary embodiment of the present invention, the core of the RITE includes the DCR which has two compression stages and a turbine, the centrifugal compressor which has one compression stage and a turbine, and two combustors (one main combustor and one ITB). The RITE core operates like two turbine engines in series, and it can be controlled that way. Numerous turbofan/turboshaft configurations can be built off the RITE core.

"In accordance with one aspect of the present invention, there is provided a gas turbine engine including a first combustion chamber, a dual compression rotor positioned behind (towards the rear) the combustion chamber, and a centrifugal compression rotor positioned behind (towards the back) the dual compression rotor. The dual compression rotor may include a first axial compression stage for compressing air in a first axial direction and a second axial compression stage for compressing the air in a second axial direction. The second axial direction is generally opposite the first axial direction.

"The first axial compression stage may include a plurality of first blades for compressing the air in the first direction, while the second axial compression stage may include a plurality of second blades for compressing the air in the second direction. Moreover, the dual compression rotor and the centrifugal compression rotor may include a plurality of tip turbine blades. The engine may also include a second combustion chamber located between the plurality of tip turbine blades of the dual compression rotor and the plurality of tip turbine blades of the centrifugal compression rotor. The second combustion chamber may be an inter-turbine burner, and the first combustion chamber may be annular shaped.

"The gas turbine engine may further include a duct positioned between the dual compression rotor and the centrifugal compression rotor. The duct may be configured for redirecting radially flowing compressed air from the centrifugal compression rotor to the second axial direction.

"In accordance with another aspect of the present invention, there is provided a gas turbine engine including a dual compression rotor supported by and rotatable about a support structure, and a centrifugal compression rotor positioned behind the dual compression rotor. The centrifugal compression rotor is supported by and rotatable about the support structure. Also, the dual compression rotor and centrifugal compression rotor may be independently rotatable relative to each other.

"The dual compression rotor may include a plurality of first axial compression blades for compressing air in a first axial direction and a plurality of second axial compression blades for compressing air in a second axial direction. The second axial direction is generally opposite the first axial direction.

"The engine may also include an annular combustion chamber positioned ahead of (towards the front) the dual compression rotor. Exhaust gas from the annular combustion chamber rotates the dual compression rotor. The engine further includes an inter-turbine burner positioned between the dual compression rotor and the centrifugal compression rotor. Exhaust gas from the inter-turbine burner rotates the centrifugal compression rotor.

"In a related aspect of the invention, the dual compression rotor and centrifugal compression rotor rotate in opposite directions, or alternatively, they may rotate in the same direction."

For the URL and additional information on this patent, see: Dale, Mark R.. Gas Turbine Engine with Dual Compression Rotor. U.S. Patent Number 8726635, filed December 17, 2012, and published online on May 20, 2014. Patent URL: http://patft.uspto.gov/netacgi/nph-Parser?Sect1=PTO2&Sect2=HITOFF&p=137&u=%2Fnetahtml%2FPTO%2Fsearch-bool.html&r=6834&f=G&l=50&co1=AND&d=PTXT&s1=20140520.PD.&OS=ISD/20140520&RS=ISD/20140520

Keywords for this news article include: Technology, The United States of America as represented by the Secretary of the Air Force.

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Source: Journal of Technology


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