Patent Description:
A gas turbine engine typically mixes a carbon based fuel with air within a combustor where it is ignited to generate a high-energy exhaust gas flow. Improving engine operating efficiencies are driven by economic and environmental demands. Most engine inefficiencies are due to heat lost when the high-energy exhaust gas flow exits the turbine and vents to atmosphere. Capture of waste heat may increase overall engine operating efficiencies.

Turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies.

<CIT> relates to a system for energy conversion including a closed cycle engine having a piston body defining a hot side and a cold side and having a piston assembly movable within the piston body.

A propulsion system for an aircraft according to one aspect of the present invention is as claimed in claim <NUM>.

A method of operating a propulsion system according to another aspect of the present invention is as claimed in claim <NUM>.

<FIG> schematically illustrates an example propulsion system <NUM> that includes at least one gas turbine engine <NUM>, a first bottoming cycle <NUM> and a second bottoming cycle <NUM> controlled to efficiently adapt waste heat power generation to operating conditions.

The example gas turbine engine <NUM> includes an optional fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. In embodiments where the engine does not directly drive a fan, the power produced may be used to drive any mechanical or electrical system of interest. If present, the fan section <NUM> drives air along a bypass flow path B. The compressor section <NUM> draws air in along a core flow path C where air is compressed and communicated to a combustor section <NUM>. In the combustor section <NUM>, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section <NUM> where energy is extracted and utilized to drive either the fan section <NUM>, or other power consuming system, and the compressor section <NUM>.

Although the disclosed non-limiting embodiment depicts a two-spool turbofan gas turbine engine, 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; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. Moreover, the features and embodiments presented are applicable to land based turbine engines.

The example engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via several bearing systems <NUM>. It should be understood that various bearing systems <NUM> at various locations may alternatively or additionally be provided.

The low speed spool <NUM> generally includes an inner shaft <NUM> that provides shaft power to drive the fan section <NUM> or a power system generating system. In one disclosed example, the inner shaft <NUM> connects a low pressure (or first) compressor section <NUM> to a low pressure (or first) turbine section <NUM>. The inner shaft <NUM> drives the fan section <NUM> through a speed change device, such as a geared architecture <NUM>, to drive the fan <NUM> (or power system) at a lower speed than the low speed spool <NUM>. The highspeed spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure (or second) compressor section <NUM> and a high pressure (or second) turbine section <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the bearing systems <NUM> about the engine central longitudinal axis A.

A combustor <NUM> is arranged between the high pressure compressor <NUM> and the high pressure turbine <NUM>. In one example, the high pressure turbine <NUM> includes at least two stages to provide a double stage high pressure turbine <NUM>. In another example, the high pressure turbine <NUM> includes only a single stage. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine.

The example low pressure turbine <NUM> has a pressure ratio that is greater than about <NUM>. The pressure ratio of the example low pressure turbine <NUM> is measured prior to an inlet of the low pressure turbine <NUM> as related to the pressure measured at the outlet of the low pressure turbine <NUM> prior to an exhaust nozzle.

The mid-turbine frame <NUM> further supports bearing systems <NUM> in the turbine section <NUM> as well as setting airflow entering the low pressure turbine <NUM>.

Airflow through the core airflow path C is compressed by the low pressure compressor <NUM> then by the high pressure compressor <NUM> mixed with fuel and ignited in the combustor <NUM> to produce high speed exhaust gases that are then expanded through the high pressure turbine <NUM> and low pressure turbine <NUM>. The mid-turbine frame <NUM> includes vanes <NUM>, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine <NUM>. Utilizing the vane <NUM> of the mid-turbine frame <NUM> as the inlet guide vane for low pressure turbine <NUM> decreases the length of the low pressure turbine <NUM> without increasing the axial length of the mid-turbine frame <NUM>. Reducing or eliminating the number of vanes in the low pressure turbine <NUM> shortens the axial length of the turbine section <NUM>. Thus, the compactness of the gas turbine engine <NUM> is increased and a higher power density may be achieved.

The disclosed gas turbine engine <NUM> in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine <NUM> includes a bypass ratio greater than about six (<NUM>), with an example embodiment being greater than about ten (<NUM>). The example geared architecture <NUM> is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about <NUM>.

In one disclosed embodiment, the gas turbine engine <NUM> includes a bypass ratio greater than about ten (<NUM>:<NUM>) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor <NUM>. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.

The fan section <NUM> of the engine <NUM> is designed for a particular flight condition -- typically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (<NUM>,<NUM>). The flight condition of <NUM> Mach and <NUM>,<NUM> ft. (<NUM>,<NUM>), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point.

In another non-limiting embodiment, the low fan pressure ratio is less than about <NUM>.

"Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(<NUM>°R)]<NUM>. The "Low corrected fan tip speed", as disclosed herein according to one non-limiting embodiment, is less than about <NUM> ft/second (<NUM>/s).

For engine embodiments that include the fan section, the fan section <NUM> comprises in one non-limiting embodiment less than about <NUM> fan blades <NUM>. In another non-limiting embodiment, the fan section <NUM> includes less than about <NUM> fan blades <NUM>. Moreover, in one disclosed embodiment the low pressure turbine <NUM> includes no more than about <NUM> turbine rotors schematically indicated at <NUM>. In another non-limiting example embodiment, the low pressure turbine <NUM> includes about <NUM> turbine rotors. A ratio between the number of fan blades <NUM> and the number of low pressure turbine rotors is between about <NUM> and about <NUM>. The example low pressure turbine <NUM> provides the driving power to rotate the fan section <NUM> and therefore the relationship between the number of turbine rotors <NUM> in the low pressure turbine <NUM> and the number of blades <NUM> in the fan section <NUM> disclose an example gas turbine engine <NUM> with increased power transfer efficiency.

Thermal energy produced through the combustion process is wasted as the high energy exhaust gas flow is vented to atmosphere after expansion through the turbine section <NUM>. The thermal energy vented to atmosphere can be used to drive other systems to produce electricity. The disclosed example propulsion system includes the first and second bottoming systems <NUM>, <NUM> that utilizes a bottoming working fluid circulated through a heat exchanger <NUM> to accept thermal energy from the high energy exhaust gas flow <NUM>. A cool bottoming working fluid <NUM> accepts thermal energy from the high energy exhaust gas flow <NUM> such that heated working fluid <NUM> is utilized by the bottoming systems <NUM>, <NUM> to produce power.

The engine <NUM> uses inlet air as the working fluid and therefore naturally adapts to power output to air density. More power, and thus, more waste heat, is produced at lower altitudes with denser air than at higher altitudes with less dense air. The example bottoming cycles <NUM> and <NUM> are closed systems and therefore do not naturally adapt to operating conditions. The individual components must remain within a relatively narrow range of operating conditions in order to function as desired. Accordingly, the operability and efficiency of the two bottoming cycle systems <NUM> and <NUM> is significantly compromised at one condition or the other when designed for a particular rate of waste heat input. The example propulsion system <NUM> provides for adaptation of the bottoming cycle systems <NUM>, <NUM> to engine operating conditions to maintain a desired operating efficiency.

Referring to <FIG>, the example propulsion system <NUM> includes first and second engine assemblies 20A and 20B that supply the first and second bottoming systems <NUM>, <NUM>. As appreciated, although two engine assemblies are disclosed by way of example, any number of engines <NUM> could be used including a single engine, four engines or any multiple of engines <NUM> within the contemplation of this disclosure. Moreover, although two bottoming cycle systems are shown and described by way of example, more than two bottoming cycle systems could effectively be implemented within the scope and contemplation of this disclosure. Additionally, although the disclosed first and second bottoming cycle systems <NUM>, <NUM> are shown and described as identically configured and sized, differently configured and sized bottoming cycle systems could be utilized and combined to provide the desired power generating compacity.

The example first and second bottoming systems <NUM>, <NUM> are a closed flow circuit of supercritical carbon dioxide (sCO2) bottoming working fluid that accepts thermal energy from the high energy exhaust gas flow <NUM> to produce a power output. In this disclosed example, the heated sCO2 working fluid flow <NUM> is expanded through a bottoming turbine 78A, 78B to generate shaft power used to drive a generator 82A, 82B and bottoming compressors 80A, 80B. The generators 82A, 82B produce electric power that can be used for functions of the aircraft and propulsion system. As appreciated, although the power recovered through expansion of the heated bottoming working fluid flow <NUM> through the turbines 78A-B are disclosed by way of example, other power recovering configurations are within the scope and contemplation of this disclosure.

The thermal energy available during engine operation differs depending on engine operating conditions. At lower altitudes, with denser warmer air, a sufficient amount of thermal energy is recoverable to operate both bottoming cycles at the optimal temperature entering turbines 78A-B, the optimal temperature entering compressors 80A-B, and the optimal flow rate for desired performance of both turbines 78A-B and compressors 80A-B. However, at higher altitudes, with thinner, air, the amount of thermal energy available for recapture may be less than desired to effectively operate both bottoming cycles. The example propulsion system <NUM> provides for adapting operation of the bottoming cycle systems <NUM>, <NUM> to operate at least one of the bottoming cycles at the desires temperatures and flow rate.

In one disclosed embodiment, the architecture is used to adjust a mass flow rate of working fluid flow through the components such that the desired temperatures and flow rates are achieved through turbines 78A-B and compressors 80A-B.

The desired mass flow of working fluid flow is provided by decelerating the working fluid flow the system, such that the mass flow rate is at a desired level for a single turbine-compressor pair 78A and 80A, enabling these to remain operable and operate at or near desired efficiency.

The reduction in flow rate results in a reduction in working fluid flow rates arriving at the waste heat exchangers 70A and 70B, enabling the reduced quantity of waste heat available at each of these heat exchangers to raise the working fluid flow rate temperature near to that of the exhaust stream. Raising the working fluid flow temperature as much as possible in these heat exchangers raises the efficiency of heat to power conversion. The working fluid flows are then combined before expansion in the turbine 78A.

In the example embodiment, the flow is split again after turbine 78A to travel through recuperators 84A and 84B and heat rejection heat exchangers 86A and 86B. Subsequently, the working fluid flow is recombined for entry into compressor 80A, enabling this compressor to operate at its optimal flow rate for high efficiency. The non-power producing compressor-turbine pair 78B and 80B is then turned off until such time as sufficient thermal energy is available from the engines for both to operate at a desired efficiency. Note that in a similar embodiment, the working fluid flow in this operating condition is not split after the turbine 78A, thus directing the full flow through recuperator 84A and cooling heat exchanger 86A. Thus, recuperator 84B and cooling heat exchanger 86B may be inactivated in tandem with the compressor-turbine pair 78B and 80B.

In the illustrated example, the bottoming working fluid is maintained at or above a supercritical point during the heat addition and turbine expansion phases of the working cycle. Due to being maintained at or above the supercritical point, the bottoming cycle systems <NUM> and <NUM> are referred to as a supercritical CO2 cycle (sCO2 cycle). The bottoming working fluid may or may not be in a supercritical state during the heat rejection and compression phases of the cycle.

Each of the example bottoming cycle systems <NUM>, <NUM> includes a corresponding bottoming turbine 78A, 78B that is coupled to a bottoming compressor 80A, 80B. The bottoming turbine 78A, 78B is coupled to drive a mechanical machine requiring shaft power, or an electric machine such as generator 82A and 82B to produce electric power.

A heat exchanger 70A, 70B are configured to transfer thermal energy from the exhaust gases <NUM> into the sCO2 working fluid flow <NUM>. The heated sCO2 working fluid flow <NUM> is expanded through a corresponding bottoming turbine 78A, 78B to produce shaft power to drive corresponding generators 82A, 82B.

The bottoming working fluid flow <NUM> exhausted from the corresponding turbine 78A, 78B is directed through a corresponding recuperator 84A, 84B that is in thermal contact with a previously cooled portion of the bottoming working fluid flow. The recuperators 84A, 84B are heat exchangers that are configured to place different temperature bottoming working fluid flow into thermal contact for heating the CO2 flow from the compressors prior to further heating in exhaust heat exchangers 70A-B. The bottoming working fluid flow from the recuperator 84A, 84B is further cooled in a corresponding ram air heat exchanger 86A, 86B. From the ram air heat exchangers 86A, 86B, the bottoming working fluid flow is compressed in corresponding bottoming compressor 80A, 80B.

The bottoming turbine 78A, 78B may expand the bottoming working fluid flow below the critical pressure immediately prior to the beginning of the working fluid cycle. This expansion is referred to as overexpansion.

The compressed bottoming working fluid flow <NUM> from the compressor 80A, 80B is then routed through the recuperator 84A, 84B and back to the heat exchanger 70A, 70B to accept heat and begin the cycle over again.

During operation in the configuration illustrated in <FIG>, both bottoming cycle systems <NUM>, <NUM> operate continuously, and similarly to the degree that engines 20A and 20B are similar. In such operation, engines 20A and 20B are generating sufficient thermal energy to efficiently operate both bottoming cycle systems <NUM>, <NUM>.

The two bottoming cycle systems <NUM>, <NUM> are coupled together by circuits that distribute bottoming working fluid flow. In this disclosed example, a first valve <NUM> splits the bottoming working fluid flow between the bottoming turbines 78A, 78B. A second valve <NUM> controls the mixing of bottoming working fluid flow exhausted from the compressors 80A, 80B. A third valve <NUM> controls the mixing of bottoming working fluid flow from the heat exchangers 70A, 70B. A controller <NUM> is provided to govern operation of the valves <NUM>, <NUM> and <NUM> based in operating condition information received from the engines 20A, 20B as well as other sources of information present on an aircraft that are indictive of operation. The controller <NUM> uses the gathered information to actuate the appropriate control valves <NUM>, <NUM>, <NUM> to isolate one of the bottoming cycle systems <NUM>, <NUM> to provide a desired efficiency.

<FIG> illustrates the propulsion system <NUM> in a take-off or low altitude configuration. In this disclosure, the low altitude is an altitude where an air density is such that the engine <NUM> generates an exhaust gas flow of a temperature above predefined desired temperature. In one example embodiment, the predefined desired temperature is that temperature that generates sufficient thermal energy to operate both bottoming cycles <NUM>, <NUM> at a desired operating condition. Moreover, in this disclosure, one example take-off configuration refers to an air speed of about <NUM> Mach at sea level, with engines operating at or near full power and the highest allowable combustion temperature. In the take-off or low altitude configuration both the first and second bottoming cycle systems <NUM>, <NUM> operate to reclaim thermal energy and produce power for use by the propulsion system <NUM> and/or aircraft systems. Both bottoming cycle systems <NUM>, <NUM> are illustrated as producing electric power by way of driving the generators 82A, 82B with the shaft power generated by expanding the heated bottoming working fluid flow through each turbine 78A, 78B.

Referring to <FIG>, the propulsion system <NUM> is configured for operation during cruise and/or high altitude conditions. In this disclosure, high altitude conditions are those conditions where the air density is below a predefined level that results in a reduced temperature and amount of thermal energy generated and exhausted as waste heat. The reduce amount of waste heat is that amount that is below the predefined level to support operation of both the first and second bottoming cycles <NUM>, <NUM>. In this disclosure cruise conditions refer to a condition that the engines 20A, 20B are designed to operate at a substantially steady state. One example cruise condition includes an air speed of about <NUM> Mach at about <NUM>,<NUM> feet (<NUM>) where the engines 20A, 20B are operating at the best fuel consumption. At the example cruise altitude, the air is thin, the combustor temperature is reduced, and the engines 20A, 20B do not generate thermal energy sufficient to drive both bottoming cycle systems <NUM>, <NUM> at desired efficiencies.

Accordingly, the controller <NUM> adapts operation by actuating the first valve <NUM> to direct the heated bottoming working fluid flow only to the bottoming turbine 78A of the first bottoming cycle system <NUM>. The controller <NUM> further actuates the second valve <NUM> to close off flow to the bottoming compressor 80B of the second bottoming cycle system <NUM>.

In the second operating condition configuration illustrated in <FIG>, heated bottoming working fluid flow from both heat exchangers 70A and 70B are directed to the one bottoming turbine 78A. Thermal energy from both engines 20A and 20B is therefore combined to provide more thermal energy than would be available from each engine 20A, 20B alone at the second operating condition. The bottoming turbine 78A therefore provides electric power and operates at a desired efficiency. In this example, the desired efficiency is an efficiency that is greater than an efficiency achievable if both turbines 78A, 78B where operating and splitting the available thermal energy at the second engine operating condition.

The bottoming working fluid flows exhausted from the turbine 78A are routed through both recuperators 84A, 84B and ram heat exchangers 86A, 86B and back to respective exhaust gas heat exchangers 70A, 70B. The bottoming working fluid flows are routed only through the bottoming compressor 80A of the first bottoming cycle system <NUM>. Operation continues according to the configuration illustrated in <FIG> until operating conditions change and the controller can actuate the first and second valves <NUM>, <NUM> to reengage the dormant second bottoming cycle system <NUM>.

The disclosed example propulsion system is adaptable to provide efficient and controlled waste heat power generation during varied operating conditions.

While described above in conjunction with a geared turbofan engine, it is appreciated that the waste heat recovery system described herein can be utilized in conjunction with any other type of turbine engine with only minor modifications that are achievable by one of skill in the art.

Claim 1:
A propulsion system (<NUM>) for an aircraft, the propulsion system (<NUM>) comprising;
at least one core engine (<NUM>; 20A, 20B) including a core flow path in communication with a compressor section (<NUM>), combustor section (<NUM>) and a turbine section (<NUM>), wherein a high energy exhaust gas flow (<NUM>) is produced in the combustor section (<NUM>) and expanded through the turbine section (<NUM>) to drive the compressor section (<NUM>);
a first bottoming cycle system (<NUM>) including a bottoming working fluid flow (<NUM>) in thermal communication with the high energy exhaust gas flow (<NUM>) generated by the at least one core engine (<NUM>; 20A, 20B), the first bottoming cycle (<NUM>) is configured to recover power from the high energy exhaust gas flow (<NUM>) in a first engine operating condition and in a second engine operating condition; and
a second bottoming cycle system (<NUM>) including the bottoming working fluid flow (<NUM>) in thermal communication with the high energy exhaust gas flow (<NUM>), the second bottoming cycle system (<NUM>) configured to recover power from the high energy exhaust gas flow (<NUM>) in the first engine operating condition, wherein:
each of the first bottoming cycle system (<NUM>) and the second bottoming cycle system (<NUM>) includes a bottoming compressor (80A, 80B) and a bottoming generator (82A, 82B) driven by a bottoming turbine (78A, 78B);
the propulsion system (<NUM>) further includes a first valve assembly (<NUM>) controlling bottoming working fluid flow (<NUM>) to the bottoming turbine (78A, 78B) of the first bottoming cycle (<NUM>) and the second bottoming cycle (<NUM>) and a second valve assembly (<NUM>) configured to control bottoming working fluid flow (<NUM>) from the bottoming compressor (80A, 80B) of the first bottoming cycle (<NUM>) and the second bottoming cycle (<NUM>); and
the propulsion system (<NUM>) further includes a controller (<NUM>) configured to operate the first valve (<NUM>) assembly and/or the second valve (<NUM>) assembly to stop bottoming working fluid flow (<NUM>) to the second bottoming cycle system (<NUM>) in response to an indication that the high energy exhaust gas flow (<NUM>) has less thermal energy than desired to operate both of the first bottoming turbine (78A) and the second bottoming turbine (78B).