HYBRID FREE-AIR GAS TURBINE ENGINE
An alternate means of generating the power to drive the compressor section of gas turbine engines is described. Electric motors embedded within gas turbines would rotate the compressor, and receive the requisite power to perform the compression work from: 1) either a high-capacity APU generator, or 2) from a relative-wind driven turbine in a hybrid power-sharing arrangement. The electric motor provides enhanced control of N1 rpm, assuring greater responsiveness to control inputs.
Not Applicable
REFERENCE TO A SEQUENCE LISTING, A TABLE, OR A PROGRAM LISTING COMPACT DISC APPENDIXNot Applicable
BACKGROUND OF THE INVENTIONThe invention relates to gas turbine engines. In particular, it relates to a device that would be used as a prime mover for aircraft or other high-speed vehicles, where the engine creates power through the combustion of fuels mixed with air that is compressed by the engine.
Turbine-powered aircraft transport hundreds of thousands of passengers efficiently and economically every day. With the advent of high-bypass turbofan engines came very economical ratios of specific fuel consumption (SFC), plus huge thrust ratings to power ever-larger commercial jets. But, growing complexity is difficult to fabricate and maintain, as well as expensive, despite the fact that turbojets have been with us now for close to three-quarters of a century.
Turbofan engines are a compromise, trading the basic design simplicity of early turbojet engines for the higher specific thrust efficiency of high-bypass fans. By using the still prodigious pressure energy created from combustion at even less than the highest fuel burning temperatures obtainable, they gain very high mechanical energy that is effectively transmitted to bypass fans, which then accelerate larger volumes of air than a ‘plain’ turbojet, though at less significant differences in intake and exhaust velocities than the turbofan engine's core.
Inherent limitations in gas turbine engine design threaten greater output, forcing designers to adopt complex mechanical interventions to achieve higher thrust potential. The most vexing of those limitations is without a doubt the need for post-combustion power turbines to withstand extreme temperatures while undergoing equally difficult stresses from rotation at very high speeds. The battle of “required tensile and shear strength required at elevated temperatures” that must be found in materials used to fabricate hot section turbine blades, and consequent intricate cooling systems needed to keep those highly-stressed assemblies within safe operating parameters, limits the potential maximum thrust and thermodynamic efficiency of turbojet/turbofan engines.
These limitations have brought about a situation where the hot turbine sections are typically one of, if not the most costly portions of gas turbines to produce and maintain, as well as that most prone to failure. Extravagant cooling systems for hot section turbines continue to be the vanguard of research for commercial and military turbofan aviation engines. The additional weight, manufacturing complexity and maintenance requirements for high-bypass spool sections of turbofan engines makes them an extravagantly expensive compromise to get the most out of engines still limited by the fragility of hot section turbines more than by any other factor.
BRIEF SUMMARY OF THE INVENTIONWhat is seen as being needed is a gas turbine design that can inherently operate at much higher combustion temperatures, thus becoming thermodynamically more efficient. High SFC values are greatly necessary—such that are comparable to current turbofan engines or better—in light of oil prices, as well as the importance of cleaner burning of fuels with regard to ecological concerns.
Simplicity, and if at all possible, less costly technology to reduce powerplant cost is a barrier that is limiting general aviation advancement, and is a primary cause for the high acquisition and operating costs associated with aviation.
It is also desirable to incorporate:
Higher Responsiveness: Gas turbine engines have a degree of response lag that is partially overcome by modern full authority digital engine controls (FADEC), but as yet is not as immediate as pilots might require in some emergency situations.
Modularity: There is the necessity of mounting engines on an airframe (or other air-enveloped vehicle) in places where they can operate at their peak, yet be protected from damage, as far as is practical. This must be reconciled with vehicular design parameters predicated on performance, ease of construction, load carrying accommodation, maintenance access and other factors. Therefore, the freedom to distribute a complete propulsion system in partial modules remotely on an airframe (or other air-enveloped vehicle) as component modules would relieve some constraints on the accommodation of that system as a whole, where linkages between modules would not themselves be prohibitively complex.
High Reliability: Commercial aviation operations require extremely reliable equipment, given the potential for great loss of life due to any mechanical failure. An inherent design simplicity incorporating components familiar to the aviation industry as proven components and/or sub-systems, though employed in different fashions in new art, is a means to ensure continued high reliability.
Reduced Aerodynamic Drag: A design that facilitates aerodynamic shape, i.e., a smaller frontal area, compared to current turbofan engines within a given thrust class.
It can be seen that a different manner in which to power the turbocompressor sections of gas turbine engines would be useful if it overcame the limitations described above.
The specification of the engine herein is not intended to be limited to use within aircraft, nor does the mention of ‘turbocompressor’, or the examples in drawings included here intend to limit the type of compressor that may be used to only axial-flow or centrifugal types, because the new art claimed here is not for a type of compressor, but by the manner in which the compressor is driven.
Included in this disclosure are five drawings of the present invention, including variations to the basic design. These drawings depict all essential elements of the device; however, they are not intended to limit the manner in which the total system is configured within a given airframe.
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- 1.—Low-pressure compressor stage inlet area.
- 2.—A high-capacity, electronically-controlled brushless electric motor, the preferred type being asynchronous AC.
- 3.—A second electric motor driving a centrifugal high-pressure compressor stage.
- 4.—Fuel line.
- 5.—Combustor area.
- 6.—Exhaust expansion nozzle.
- 7.—Centrifugal compressor impeller disk.
- 8.—Bleed-air driven auxiliary generator.
- 9. FAT relative-wind turbine with constant speed hub.
- 10.—An intermediate gear or torque converter fluid coupling incorporating a minimum of two selectable step-up ratios.
- 11.—Electric generator, electronically controlled in the preferred embodiment.
- 12.—A power source for an independent generator, typically being a relatively small turbo-shaft engine as used in most standard aircraft auxiliary power units (APU). The preferred embodiment is for an APU that uses the same fuel as the thrust engine.
- 13.—An intermediate clutch assembly required between co-located APU and FAT modules.
- 14.—The basic embodiment of an APU module, as opposed to the front end of the depicted power train.
- 15.—The basic embodiment of a FAT module.
The invention may be visualized as being composed of three major component modules, where a conventional turbocompressor section using axial and/or centrifugal stages is still present as typical to conventional gas turbine engines.
In contradiction to conventional design, however, this invention has a compressor section that is not powered by a hot section (gas generator) power turbine mechanically linked to it by a rotating shaft, but is instead powered by one or more high-output electric motors; followed by fuel burners and then an exhaust nozzle. This portion of the engine is referred to as the ‘thrust module’. No advance over prior art is intended in the combustion or the exhaust nozzle, though it will be explained later that combustion section components must be engineered, using established technology, for its stationary parts to accommodate much hotter ignited gas flows.
The turbocompressor configuration described above needs a source of electrical energy to power the compression work. In U.S. Pat. Nos. 3,527,055 and 3,705,775, the source of electrical energy perceived by those inventors included magnetoplasmadynamic flux and/or conventional electric generators connected to hot power turbines. Unfortunately, these two ideas neither solved the limitations mentioned earlier, nor ameliorated them. Driving a generator by a hot section power turbine continues the problem of turbine component stress at high centrifugal-centripetal loads and temperatures, and adds to that the susceptibility of generators to excessive heat if placed internally. If magnetoplasmadynamic flux were used to generate electrical current, it would necessarily force gas turbine engines into a much higher plateau of heat stress (>5000° F.) where even many non-rotating parts would be marginalized, especially if subjected to regular service at such extremes; there are also limitations as to how much current can realistically be produced by that method.
Contemporary electric motors of both brushless DC and asynchronous induction AC types plus the use of exotic materials for magnets now offer very-high power-to-weight ratios, as well as electronic controls and other features that are favorable to aviation needs. They are suited to driving a turbocompressor, provided that sufficient electrical energy can be generated to support the air compression task.
Behind that is the exhaust nozzle, lacking any power turbine components, and of a general shape allowing expansion of hot gases exiting the burners at much higher than the temperatures found in most contemporary gas turbine engines, though still far below the theoretical magnetoplasmadynamic flow range. Typical turbofan parameters currently limit hot turbine inlet temperatures to between 1,200°-2,400° Fahreinheit; yet this new engine and its minimal and mostly fixed exhaust nozzle components could safely achieve exhaust temperatures much closer to the hydrocarbon stoichiometric combustion upper limit (around 4,200° Fahreinheit under elevated pressure). This is because it lacks a power turbine section, so the exhaust nozzle, in basic shape and design, becomes cross-sectionally rocket nozzle of largely fixed components subjected to relatively modest mechanical forces. Without the rotating parts of a power generating turbine stage, the exhaust section can thus be allowed to entertain much higher temperatures to optimize fuel efficiency and thrust; 3,500-3,700° Fahreinheit, or higher, in ‘exhaust nozzle inlet temperatures’ are thus within achievable limits. Being also free of the constrictions to jet flow caused by hot section turbine apparatus, the exhaust nozzle is only required to facilitate expansion of the exhaust gases, thus rendering maximized jet exhaust flow. Therefore, this hybrid engine offers quite high thermodynamic efficiency via higher fuel-air burning temperatures that are not limited by turbine blade technology. Greater overall thermodynamic efficiency is also helped because the hybrid engine lacks the substantial heat loss associated with turbine blade cooling systems.
The primary electrical power producer in-flight becomes, then, a ‘free-air’, windmilling turbine (i.e., ‘free-air turbine’, or “FAT”). In appearance, it is not unlike an aviation propeller. It is, however, external to the engine, exposed to the relative wind flow over the aircraft or other high-speed vehicle while in forward motion. It could be likened to a ground-based, fixed, wind-powered generator; however, it would operate in a much higher speed regime at aircraft flight speeds, and might feature wider chord blades to increase the disc solidity factor and thereby augment its energy capturing potential. Because of the forward speeds involved, the energy that can be derived at even relatively low aircraft speeds, versus the power needed to compress air for the thrust engine, quickly converge into compatible values. In fact, above airspeeds of about 100 knots per hour, the energy can be significant, and at an attractively small aerodynamic drag penalty. A conventional aviation constant-speed propeller hub assembly maintains the wind turbine's rotational speed within a desired range, matching its speed to a coupled generator. This relative wind-turbine, with its associated generator, form the ‘FAT module’.
It might be noted that, above about 300 knots, a modest overdrive gear or continuously variable transmission between the constant-speed hub and generator would increase the efficiency of the wind-turbine output by allowing for high FAT blade tip speeds at lower forward velocities, but facilitate lower tip speeds at higher forward velocities. The relative wind at high forward velocities can impart such high torsional rigidity on a wind-driven turbine that very high rotational speeds are not necessary to power a generator at the required power inputs. Thus, high FAT tip velocities approaching transonic values can be avoided that would otherwise compromise wind-turbine efficiency at high subsonic forward velocities. As conventional propellers approach forward velocities at or above Mach 0.7, the forward vector component of their high rotational speeds plus the aircraft forward velocity puts blade tips in the transonic range long before the forward velocity of the aircraft reaches that threshold where airfoil efficiency drops and other problems arise. The high rotational speeds of propellers on the most powerful propeller planes are predicated by their having to impart energy onto air and accelerate it. This is fundamentally different from a turbine that receives energy from a relative wind. Though aerodynamic principles mutually apply to both propellers and turbines, the net difference is that the propeller must turn fast to translate energy, whereas the turbine can receive a great deal of energy at little rotational speed; its torque is in fact is potentially highest at zero rpm at any given forward speed.
Thus, the gearing stage would multiply the high-torque, low-rpm rotational speeds of the FAT (at high forward velocities) to a higher rpm value acceptable for the generator. Alternatively, electronic controls on the generator could in some cases achieve a favorable output without resorting to a geared coupling.
Conventional parts can be used to form most of the components for the ‘FAT module’ in small- to medium-sized aircraft applications. Of course, without the need for mechanical couplings between a generator and the turbocompressor's electric motor, there is no need to reconcile differing rotational speeds between them. This would be governed by electronic motor and generator control circuitry.
At rest, to start this new engine, a supplementary generator is needed that could be adequately supplied by conventional onboard aircraft auxiliary power units (APUs), though in almost all cases, significantly larger outputs would be desired in a given ‘APU module’ installation than what would normally be installed in a given airframe today. Nevertheless, there already exist highly compact turboshaft engines coupled to high-capacity generators to support the invention described here, and in sizes that would equate to most thrust categories in commercial aviation applications. This however, may be viewed only as a starting and initial operation sequence prior to full operation via the FAT module. Being able to power the turbocompressor by either the FAT or APU modules is the reason for the ‘hybrid’ appellation.
In this manner, this ‘hybrid engine’ would, in typical power categories, not be appreciably heavier than a conventional turbofan engine of comparable output (at speed), because a conventional turbofan engine's hot power turbine section, plus the large high-bypass fan section it must drive, including ancillary mechanical couplings, bearings, lubrication and cooling systems, etc., are characteristically heavy, too, when compared to the simple bell structure of a fixed nozzle exhaust on this invention. The conventional bypass section and power turbine components in a turbofan offset the weight penalty of the ‘FAT module’. It is possible, especially in large thrust categories, that a ‘hybrid engine’ could be lighter overall than a conventional turbofan engine of comparable thrust because of the size of bypass fan sections used conventionally. (Such comparison must subtract the weight of APU installations in a conventional turbofan-powered aircraft from the weight penalty of a larger ‘APU module’ with the hybrid engine.)
However, to achieve equivalent thrust in such a comparison, a larger turbocompressor mass flow would be required in the hybrid engine than that of the turbofan engine core; the invention would not need to match the overall mass flow of the turbofan, but would require sufficient increase (over the turbofan core flow) in the hybrid engine's throughput to match the turbofan's thrust after the hybrid engine's much higher operating temperatures are applied. Given a simple analogy of only a 30 percent higher mass flow (over an equivalent turbofan engine core throughput) with a nearly two-fold increase in operating temperature in some cases—given thermodynamically exponential gain—to achieve competitively similar thrust to a comparable turbofan, it easy to see that the hybrid engine can be smaller overall, cheaper and easier to build, while offering high thrust performance. Thermodynamic principles imply that in such a scenario, a hybrid engine could be more economical than an ‘equivalent’ turbofan.
There is the proviso to this concept that practical and economically-satisfactory APU installations on any size of aircraft would likely limit the highest static thrust at the point of brake release for takeoff available with a hybrid engine to less than that available with current comparable turbofan engines. This would be due to the amount of power required to operate a turbocompressor at high static mass flows. Thus, such performance would be dependent upon the power rating of the APU selected, and so, practically speaking, a size-limited and cost-effective choice of APU could render a situation where a hybrid engine's static thrust could be significantly less than that of a given turbofan, though they might otherwise be comparable (or the hybrid engine even superior) in thrust performance during climb, cruise and descent flight phases. This weakness in the hybrid is, however, made less difficult by the fact that at typical jet aircraft takeoff decision (V1) or rotation (Vr) speeds of 110-160 knots or so, the hybrid engine's FAT module could be made to supply 20-40% or more of the total current available at rotation, making climb performance thereafter as good as a turbofan certainly possible.
FAT modules could be facilitated toward early production of meaningful current by briefly being spun-up prior to brake release and the beginning of the takeoff run. The rotational speeds need not be high; this would be intended to help overcome the FAT unit's inertia, if at rest. In a design configuration that co-locates the thrust module with the FAT module ahead of the turbocompressor, a small amount of auto-rotation might be induced by the air impelled into the turbocompressor (see
It is also true that short-duration, ultra-high power output features can be designed into APUs using conventional technologies to facilitate takeoff needs at reasonable cost. Such short-duration performance enhancement was addressed by U.S. Pat. No. 4,759,178 and others, and can be found in many technologies already marketed. That could then more adequately supply hybrid engine needs.
When airborne and accelerating in climb, wind-turbine power output increases rapidly with the forward velocity, while the power needed to turn the turbocompressor at a maximum output falls just as dramatically with higher altitudes. Above approximately 220 knots or so (a typical climb speed for turbine-powered aircraft), a ‘FAT module’ could be expected to produce all of the power needed for the turbocompressor, alone. It would then remain so throughout the rest of the flight. On commercial flights with high onboard power requirements, the APU could be powered down as the FAT module takes over, but continue operation to meet in-flight electrical loads; on general aviation aircraft with lesser in-flight power needs, the APU would likely be shut down until precautionarily restarted during approach to ensure that rapid go-around power were available, if needed.
In a typical flight profile, the APU would be required to operate during the takeoff and early climb phases for only about fifteen minutes or less, and at maximum output for likely less than five minutes of that (not including ground operations). Obviously, as an operational safety backup, the APU would be required to have high-altitude restart capability, as found in many APU units on the market today, and addressed by patents such as U.S. Pat. No. 7,089,744.
It is estimated that the cost of the hybrid engine, including a relatively large APU, could be kept at 30-40% cheaper than a comparable turbofan because of the simpler and less costly components employed. A significant proportion of the total hybrid engine acquisition cost would actually rest in the APU.
There are two other components to this hybrid engine to make it self-contained. The first of those would be a stored energy (SE) module, i.e., a battery pack, fuel cell or similar; of course, aircraft generally have batteries, too. Given the current rate of advance in stored energy technology for contemporary alternative energy programs and energy conservation mandates, perhaps a lightweight stored energy module may become available later on that could add a significant portion to the total electrical load needed for takeoff, partially relieving the load required from an APU. At this stage of development, however, the ‘SE module’ requirement relative to the hybrid engine relates only to APU starting and is thus not a major issue, and not a unique component.
The last of the ancillary devices for self-contained operation would be a much smaller, engine-driven generator (or alternator) that in a preferred configuration would be one integral to the thrust module and attached to its casing, getting its power from an engine bleed-air driven centrifugal turbine. Such a small, simple device would only be intended to support battery charging (and other aircraft power needs) when and if the APU module is turned off—or fails—but could be designed to add substantially to the onboard power grid.
With three generating sources—the FAT module, APU module and thrust module alternator—this engine design offers intrinsically high reliability, simplicity, simpler manufacture and potentially significant lower acquisition cost, along with the promise of competitively high fuel economy. The electrically-driven offers a side benefit of quicker response to control inputs, because the electric motor determines N1 (compressor rpm) speed quite quickly without waiting for a reaction to increased fuel burn. In the same way, compressor stalls can be more proactively intervened.
The generator driven by the APU could be installed in such a way as to be the same unit used by the ‘FAT module’ to save weight, given the addition of clutches to engage or disengage one the other driver when not needed. That is shown in
Since the APU and FAT modules do not need to be mechanically connected to the thrust engine, they can be either co-located in a common nacelle, compartment or bay, or distributed separately at various points in an airframe (or other vehicle) to offer the greatest design flexibility. Laying power cables is of course a lighter and easier proposition than rigging mechanical couplings. Thus, the thrust module might be mounted conventionally in an underwing nacelle or aft fuselage pod, etc., while the associated FAT module could be mounted within a tail cone or on tail empennage fins (the preferred location for aerodynamic stability and esthetic reasons—see FIG. 4)), as well as at wing leading or trailing edge locations . . . anywhere there is a relatively undisturbed airflow available for it operate smoothly. The configuration depicted in
The SE module is typically located where its location facilitates maintenance and is conducive to weight and balance considerations.
Not co-locating primary components offers a bonus to flight safety; physical damage to one module is less likely to extend to other modules. There exists the possibility, too, of using shared support modules to serve multiple thrust-engine modules; ergo, one FAT module supplying in-flight power to two thrust modules, etc.
Claims
1. A windmilling, free-air turbine external to a gas turbine engine powerplant that powers the compressor of that engine via energy taken from the relative wind passing over a vehicle in motion.
- a.) where the turbine in (1) turns an electrical generator to produce electrical current to power the turbo-compressor; or,
- b.) where the turbine in (1) produces the energy to power the turbocompressor, rotating it at desired operating speeds via mechanical couplings and/or clutches.
Type: Application
Filed: Nov 19, 2010
Publication Date: May 24, 2012
Inventor: Rudolph Allen SHELLEY (BEAVERTON, OR)
Application Number: 12/950,859
International Classification: F01D 15/08 (20060101); F01D 15/10 (20060101);
