EP1977082A2 - Soufflante de dérivation et compresseur de turbine électrique pour propulsion hybride - Google Patents

Soufflante de dérivation et compresseur de turbine électrique pour propulsion hybride

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Publication number
EP1977082A2
EP1977082A2 EP07716374A EP07716374A EP1977082A2 EP 1977082 A2 EP1977082 A2 EP 1977082A2 EP 07716374 A EP07716374 A EP 07716374A EP 07716374 A EP07716374 A EP 07716374A EP 1977082 A2 EP1977082 A2 EP 1977082A2
Authority
EP
European Patent Office
Prior art keywords
compressor
turbine
flow
fan
stage
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP07716374A
Other languages
German (de)
English (en)
Inventor
Richard H. Lugg
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Sonic Blue Aerospace Inc
Original Assignee
Sonic Blue Aerospace Inc
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Filing date
Publication date
Application filed by Sonic Blue Aerospace Inc filed Critical Sonic Blue Aerospace Inc
Publication of EP1977082A2 publication Critical patent/EP1977082A2/fr
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D15/00—Adaptations of machines or engines for special use; Combinations of engines with devices driven thereby
    • F01D15/10—Adaptations for driving, or combinations with, electric generators
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D19/00—Axial-flow pumps
    • F04D19/02—Multi-stage pumps
    • F04D19/026—Multi-stage pumps with a plurality of shafts rotating at different speeds
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D25/00—Pumping installations or systems
    • F04D25/02—Units comprising pumps and their driving means
    • F04D25/06—Units comprising pumps and their driving means the pump being electrically driven
    • F04D25/0606—Units comprising pumps and their driving means the pump being electrically driven the electric motor being specially adapted for integration in the pump
    • F04D25/0613—Units comprising pumps and their driving means the pump being electrically driven the electric motor being specially adapted for integration in the pump the electric motor being of the inside-out type, i.e. the rotor is arranged radially outside a central stator
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the invention relates to an electric turbine bypass fan and compressor for hybrid propulsion.
  • thermodynamic cycle scheme i.e. the thermodynamic relations of cycle media in the process of power production.
  • thermodynamic cycle scheme i.e. the thermodynamic relations of cycle media in the process of power production.
  • it involves the introduction of fuel heat input at maximum possible temperature, compression and expansion, at maximum compressor and turbine efficiency, along with the release of non- convertibte heat to ambient temperature at minimum loss.
  • Gas turbine engines, and the devices that are powered by gas turbine engines, are limited in overall design and performance by mechanical, material, and thermodynamic laws. They are further constricted by the design limitations of the three elements that make up the baseline design of gas turbine engines: the compressor, the combustor and the turbine. In turbines for aircraft, these three engine sections are contained inside of the outer turbine casing and are centered on a load bearing drive shaft that connects the turbine (on the rearward portion of the drive shaft) with the compressor (on the forward portion of the drive shaft).
  • the drive shaft is a twin or triple spool design, consisting of two or three concentric rotating shafts nested one inside the other.
  • the different spools allow the turbine assembly and the compressor assembly, each of which is connected to one of the spools of the drive shaft, to rotate at different speeds: the turbine is optimized to run at one particular speed for combustion and thrust processes, and the compressor is optimized at a different speed to more efficiently compress incoming air at the inlet face.
  • the difference in speeds of the spools is typically accomplished by reduction gears.
  • the compressor assembly consists of several compressor stages, each of which is made up of a rotor and a diffuser.
  • the rotor is a series of rotating airfoil blades, or fans (attached to the shaft), which converge the air, i.e., compressing the volume of air on the intake side of the blade into a smaller volume of air at exit.
  • Adjacent to each rotor is a diffuser.
  • the diffuser is a fixed, non-rotating disc of airfoil stators that expands the volume of the incoming high pressure air, now at higher velocity after exiting the adjacent rotor, by having the air pass from a narrow opening on the intake side of the diffuser into a gradually enlarging chamber that slows and lowers the pressure of the air.
  • Each compressor stage is made up of a compressor rotor and a diffuser disc. There are as many stages of the compressor as are required to get the air to the required air temperature and compression ratio (in high performance aircraft turbines usually in between 40:1 to 65:1 dependent on combuster design, flight and speed envelope and turbine thrust requirements) prior to entering the combustor.
  • the higher pressure and higher temperature air mixes in a swirl of hot liquid fuel and ignites to form a controllable flame front.
  • the flame front expands as it combusts, rotating and driving turbine blades as the flame front exits the engine.
  • the turbine assembly consists of several sets of rotating turbine blades connected to the drive shaft and angled so that the thrust of the flame front causes the blades to rotate.
  • the turbine blades being connected to the drive shaft, cause the drive shaft to rotate and thus the compressor blades to rotate.
  • Turbomach ⁇ nary design must be optimized in terms of flow efficiency, high temperature blade cooling methods, rotor speed, and turbine compressor driving connections on the basis of sound rotor dynamics.
  • Many technical specialties are interwoven in a design; e.g., axial flow air compressors involve the intersection of thermodynamics, aerodynamics, structures, materials, manufacturing processes, and controls.
  • selection of rotational speed is complex in current turbomachinary designs using drive shafts. It largely depends on the balance of the requirements of the three major components on the common shaft — the by-pass fan, compressor, and turbine. Because of requirements for differential compression and associated rotational speeds, the drive shaft is multi-segmented with one shaft running inside another. In an electric turbine by-pass fan and compressor system, eliminating the drive shaft leads to a more refined approach to differential staging of the fan to the compressor, and the interrelation of thermodynamics and efficiencies with interstages in multi-axial compressor designs.
  • the overall layout of multiple compression stages in turbomachines is driven by the objective of maximizing the performance of the first transonic turbine stage and its associated impact on subsequent turbine stages and their efficiences of power extraction from the combusting gases.
  • Electric turbo compressor-compounding eliminates the mechanical coupling to the engine crankshaft, thereby eliminating the need for a crankshaft forward of the combustor. This provides additional flexibility in packaging the thermodynamic cycle scheme and its design in the turbine.
  • the compressor- compounding also provides more control flexibility in that the amount of power extracted can be varied, allowing for control of engine thermodynamics, pressure ratio, fuel consumption, mass airflow, entropy and endothermic reactions and nitrogen oxide (NOX) and carbon dioxide (CO2) formation.
  • NOX nitrogen oxide
  • CO2 carbon dioxide
  • the compressor-compounding can be operated as a ring-generator with embedded systems controls for switching and generate large amounts of power for other electric payloads on an airframe.
  • the compressor of the present invention has one or more rotor stages (compressor and diffuser), each being driven by one or more electric ring motors, the compressor rotor stages are designed and tuned more precisely to the compression ratio to be attained within the turbine design operating characteristics, thrust requirements and flight envelope. This allows for optimal aerodynamic design and efficiencies of the rotor stages in the compressor and subsequently the possibility of fewer stages needed to achieve the required compression ratios for operation of the turbine. The result is a significant potential in weight savings. Because each compressor rotor may be driven independently and at different speeds, the engine may be used more efficiently at different stages of the flight envelope. The impact of the present invention, its innovation and the unique aspect it can impose on current turbomachinary layout design, thermodynamic cycles, and thermal efficiencies, which can improve power production, is dramatic.
  • An additional object of the present invention is to use the engine more efficiently at different stages of the flight envelope. Still another object of the invention is to provide conductive pathways to power the ring motor magnetics via the generator location.
  • Yet another object is to provide a novel and unique configuration of forming electrical conductive pathways in rotational turbomachinary components.
  • Another object is to reduce the number of rotor/diffuser compressor stages.
  • Figure 1 is a graph of temperature versus cycle
  • Figure 1B is a graph as depicted in Figure 1 and an illustration of the gas compression process
  • FIG. 2 is a corresponding illustration of the gas compression process
  • Figure 3 is an equation and legend of vehicle weight, magnet coil, levitation, etc.
  • Figure 3B shows the magnetic drag versus aerodynamic drag
  • Figure 4 is an equation depicting the intersection of flight condition, design, and atmospheric properties
  • FIG. 5 shows the steady and unsteady states of the invention described herein
  • Figure 6 shows the streamtube and other components as a function of the cycle shown
  • Figure 7 depicts the guide vanes, rotor and stator with corresponding notations regarding tangential velocity increase
  • Figure 8 shows the stability boundary of Tc versus m
  • Figure 9 is a depiction of the stator and rotor with notiations regarding tangential velocity
  • Figure 10 is a graph of A/A versus M
  • Figure 11 is a depiction of the preferred embodiment of the present invention.
  • the key operations of the electric by-pass fan and electric turbocompressor-compounding compressor turbine system are that they are disengaged or engaged electrically, so that combustion cycles, compressor ratios, compressor cooling, thrust, and electric generation can be arranged and optimized for high thermodynamic and combustion efficiencies across the entire flight envelope, regardless of altitude, air density, temperature and other operating constraints.
  • the electric by-pass fan and/or electric turbocompressor- compounding system is designed to operate at ideal compression, combustion and burn efficiencies, and at higher temperatures, throughout a broader range of operation, from low subsonic (Mach 0.3) to high supersonic (Mach 2.8+) flight speeds. This is due to the magnetic, thermodynamic, mechanical and electric technologies that enable electric compression and bypass fan operation.
  • the pressure ratio compressibility can be matched to multiple design point operating conditions.
  • the electric by-pass fan has one or more low- bypass fans and/or electric compressor stages making up a compound compression system.
  • the air flows into the electric by-pass fan and/or the electric compressor in an axial direction through a series of rotating rotor blades, and stationary stator vanes that are concentric with the axis of rotation.
  • the flow path in the axial electric ring by-pass fan and the electric multi compressor stages (turbocompressor-compounding system) decreases in cross-sectional area in the direction of flow. The decrease in cross- sectional area is in proportion to the increased density of the air as the compression progresses from stage to stage
  • the preferred embodiment of the present invention has one or more stages comprising a compressor and diffuser. Each stage is driven by one or more electric ring motors.
  • the compressor rotor stages are designed and tuned more precisely to the compression ratio to be attained within the turbine design operating characteristics, thrust requirements and flight envelope. They are independent form one another, which offers greater flexibility in the generation of compression, maximum pressure ratio attained, aerothermodynamic generation heating ratios, and high endothermic and entropic combustion and fuel burn oxidation optimization, ultimately being passed on in the combustion cycle to a highly efficient fuel burn. This allows specifically for optimal aerodynamic design and efficiencies of the rotor stages in the compressor, and accordingly, the possibility of fewer stages needed (hence potential significant weight savings) to achieve the required compression ratios for operation of the turbine. Because each compressor rotor can be driven independently and at different speeds, the engine may be used more efficiently at different stages of the flight envelope as a combustion turbine machine.
  • a compressor stage enables the compressor rotor to generate higher torque than a shaft driven compressor rotor (wherein the compressor fan rotors are being driven from the tip of the blade at the circumference of the rotor rather than from the root or hub, and the leverage moments required to overcome mechanical loading are in an order of magnitude less) and enables the compressor stage to optimized typically constrained design variables, including those set forth below;
  • ⁇ Chord to height ratios, C/H can be increased due to higher stage loading conditions
  • Mufti-speed stages allow for additional compression of the fluid flow direction between the rotor and the fixed stator, wherein the airfoil profile is distributed differently across the blade airfoil camber line to reduce drag tosses and raise stage coefficient efficiency;
  • Inlet flow angle can be marginalized to drag a broader chord airfoil at a shorter blade length for increased efficiency.
  • Stage numbers being reduced allows for a compressor design to effect mean-line diffusion factors "D”, mean-line solidity u @", and polytropic efficiency "E”, thus effecting the overall efficiency of the compressor machine as a compounding medium and consequently the overall compressor ratio across the machine.
  • Higher compression ratios
  • stage inlet Mach number decreases through a multi-stage compressor, and stage pressure ratios of repeating-rows in the repeating stages also decrease.
  • Mach number can be maintained or increased, hub/tip ratio reduced, axial Mach number increased, and the total change in temperature across each stage can be raised to cause a positive effect on the atomization of the fuel as the compressed air (and thus heated air) enters the combustor;
  • Inlet guide vanes are designed to add swirl in the direction of rotor motion to lower the Mach number of the flow relative to the rotor blades.
  • the first rotor stage velocity, and angular vector are adjusted to match more closely the inlet Mach number.
  • Energy conservation is increased as the mass flow moves to the second compressor stage.
  • the second rotor stage is set at the optimum velocity to match the falling Mach number due to swirl and the velocity vector of the preceding rotor stage in the electric compressor, however, across the electric compressor energy is conserved, compression ratio raised to a higher level per each given unit of energy compared to current art of multi- axial compounding compressors using drive shafts;
  • Surge and choke lines that bind the operating range of a gas turbine engine are set for compressor aerodynamic steady state performance maximization and define the end points of operation for the compressor within the turbomachine.
  • an engine compressor is designed with a surge margin.
  • surge margins as a design point for steady performance and operation are employed due to transient conditions that move the compressor operating point (compression ratio, mass air flow, mass and stage loading, temperature rise and turbine/compressor rise ratio) close to the surge line.
  • Large surge margins place the compressor operating line and end points far from the surge line and preclude the operation at the desired peak pressure rise or maximum efficiency region of the compressor and the turbine.
  • Two types of instability can develop in a compressor; surge and stall.
  • Surge is a global asymmetric oscillation of flow through the compressor which can reverse the flow during a portion of the surge cycle.
  • Rotating stall is a local flow deficit that rotates around the compressor annulus. This flow deficit, or cell, is a region in which the local mass flow is near zero. Gas turbine engine steady performance can be optimized and improved. Rotating stall may consist of one or more multiple cells that rotate around the compressor at an angular speed which is a fraction of the rotor speed. This instability results in a loss of compressor performance that may require the shut down of the engine to clear.
  • variable speed compressor stages operates at different speeds and therefore adjust the velocity of flow, angular velocity, mach number flow and its angular vector and shock, pressure ratio and compression efficiency, so that the surge margin, or compressor stall point, is reduced and controlled. Consequently operation at peak pressure rise is maintained and the surge point is moved closer to the maximum compressor efficiency operating point without crossing it into stall or surge conditions.
  • each stage has an optimized RPM and velocity of flow Mach number set from one preceding stage to the next in the invention disclosed herein of an electric, axial flow compressor.
  • the design point of the electric compressor is set to maintain velocity and pressure of exit flow from each stator (fixed vane) of a rotor stage to the follow on rotor stage, rotating at a different RPM, but set to the optimization pressure, temperature and Mach number of the flow to maximize pressure rise between the stages.
  • the flow rate is lowered between the stages to improve the aerodynamic performance of the rotor, namely aerodynamic efficiency or stage efficiency.
  • the stage efficiency of an adiabatic multistage compressor is defined as the ratio of the ideal work per unit mass of flow to the actual work per unit mass flow between the same total pressures.
  • the other measure of efficiency which is beneficial in the preliminary design of compressors is the polytropic efficiency.
  • the polytropic efficiency of an adiabatic compressor is defined as the ratio of the ideal work per unit mass to the actual work per unit mass for a differential pressure change. In the limit, as pressure ratio approaches on for a given stage, the stage efficiency approaches the polytropic efficiency.
  • Axial flow compressors designed for jet engines in the 1980s have a polytropic efficiency of about 0.88, whereas the compressors of current art have polytropic efficiencies of about 0.90.
  • the electric mult-iaxial ringmotor compressor discussed here, baseline design on polytropic efficiency improvements come from aerodynamic drag reductions from the magnetically levitated air bearing of the compressor stages and axial hub drag reduction (discussed later in this paper), as there is no hub nor shaft.
  • Design estimates for polytropic efficiency improvements are in the range of 0.02 - 0.05, for potential improvements in the range of 0.92 - 0.95.
  • enthalpy and efficiency management cannot be done through the micro-management of the airflow between one compressor stage (rotor stage) and the next because every component is connected to a shaft.
  • the present invention is a multistage shaftless design or single stage shaftless electric compressor.
  • every stator row is a slower moving airfoil blade row, thus having the capacity to add net energy to the flow, as well as acting as a conversion device to the flow, adding kinetic energy to the flow and raising the static pressure simultaneously of the flow.
  • compressor of the present invention has one or more rotor stages, each being driven by one or more electric motors, the compressor rotor stages are designed and tuned more precisely to the compression ratio to be attained within the turbine design operating characteristics, thrust requirements and flight envelope. This allows for optimal aerodynamic design and efficiencies of the rotor stages in the compressor with fewer stages needed to achieve the required compression ratios for operation of the turbine.
  • each compressor rotor may be driven independently and at different speeds, the engine may be used more efficiently at different stages of the flight envelope.
  • each compressor rotor stage is adjacent to an electrical conductive pathway diffuser stage and can be run independently of the others with motor controllers at the outer ring of each stage.
  • This configuration of forming electrical conductive pathways in rotational turbomachinary components is also novel and unique. This configuration of the electrical compressor allows for aerodynamic optimization to meet compression ratios otherwise considered unachievable with a fixed drive shaft driven compressor.
  • the configuration of the present invention not only provides thrust as bypass air around the combu ⁇ tor but also acts as a supercharger to the turbine.
  • mass airflow is accelerated exponentially, in relation to the velocity of the air in question, at any given rate of change in time.
  • the supercharging effect upon the turbine is due to the very high optimal pressures now achievable by the electric compressor, which can be tuned to the flight condition and altitude for which the electric compressor fan is designed.
  • the preferred embodiment of the invention further comprises a gas turbine engine in which the turbine rotors and the compressor rotors are not connected by a drive shaft. Rather, the turbine rotors are connected to a drive shaft which joins them and are in turn a series of ring generators (dependent on the number of turbine disks) that transforms the mechanical energy from the turbine to electrical energy for the multi-axial ringmotor compressor and fan.
  • a compressor rotor, and a low-bypass turbofan as in a supersonic configuration is driven electrically, and not driven by the drive shaft from the turbine as in the current art; and the section of the engine that constitutes the compressor section is not connected by a drive shaft to the turbine section or the combustor section of the engine.
  • compressor stages may be optimized aerodynamically, and compressed air ratios, fractional and mass air flow flows can be optimized to each flight condition (idle, acceleration, afterburner, cruise, deceleration), maximizing the efficiencies of the compressor.
  • the electric compressor turbine engine functions as a mass-flow dynamic device, separate from the diffuser stages, combustor and turbine.
  • the electric compressor is ultimately used as a throttling and engine cycle mechanism, and its velocity is independent of the turbine engine, but contributes largely to achieving required compressor ratios for combustion, mass air flow, by-pass air for thrust, and optimal fuel burn.
  • This permits high compression ratios and finely tuned air pressures, engine cycle efficiencies independent of combustion, consistent fuel burn, effective temperature operation and cooling. Higher energy levels are achievable, and broader flight envelopes are possible because the compressor stage acts independently.
  • a further advantage of the electrically driven compressor is that rotational speed of the rotor stages does not suffer from spool up or spool down time (the time spent increasing or decreasing the rotational speed of the drive shaft) as is the case in traditional turbine designs, and the speed of the compressor rotors can be more quickly adjusted to achieve optimum performance of the engine based on different flight conditions, airframe loads, and optimal combustion performance.
  • the invention demonstrates that a multi-disc, turbofan assembly of the invention concept, because each fan disc is driven independently by an electric ring motor, the fan pressure ratio (hence the mass flow ratio) and the bypass ratio can be varied and optimized against temperature across the main components, fan, compressor and turbine.
  • Figure 1.0b is a depiction of gas turbine engine station numbering with compressor defined. The Brayton cycle depicted in Figure 1.0a is included.
  • T ⁇ & PT are more easily measured quantities than static properties (T and p).
  • Figure 2.0a depicts a schematic with appropriate component notations, compressor defined.
  • t t temperature by noting that they are related by the condition that the power used by the compressor is equal to the power extracted by the turbine.
  • This assumes an adiabatic condition of enthalpy of mass flow, temperature, and velocity across the combustor (between the compressor/fan and the turbine) and electromagnetic power consumption for the compressor ring motor drive and levitation coils is equated to with power production (including losses and power conditioning) from either turbine ring generators or MHD drive using alkaline seeded exhaust in the electric compressor concept.
  • the following step denotes the writing of an equation which represents the temperature rise across the combustor in ratio with the change in compression/change in temperature and in terms of
  • the equation represents the ideal where by in compression Delta T is minimized, and this is most accomplished with a multistage, electric ringmotor compressor, where conservation of energy is maximized, enthalpy decay is minimized by the two largest variables against degrading performance; aerodynamic drag and mechanical friction. Magnetic air bearings (Maglev) address this, and it is unique to this invention.
  • the equation follows:
  • the next step involves re-writing the equation for specific impulse, enthalpy rise, Mach number and fuel flow/heating value ratio in terms of these same parameters. This is done by beginning with writing the First Law across the combustor to relate the fuel flow rate and heating value of the fuel to the total enthalpy rise.
  • the ideal thermal efficiency is:
  • compressor area, and subsequent stage diameter design optimization is critical in defining further performance advantages as magnetic drag reduces with diameter and raise in shear pressure to achieve high energy level densities.
  • Analysis such as this can be used to define feasible pressure versus velocity profiles such as that shaded in Figure 3B.
  • This graphic relates to research on power magnetics of PRT Maglev vehicles using Halbach Array tracks for levitation and propulsion. Larger vehicles, lower magnetic drags, and different vehicle-tube clearances would change the window of opportunity.
  • thermodynamic results of compressors and turbines p's and t's.
  • thermodynamic mathematical expressions e.g. moving blades
  • Delta E 1 The amount of energy required to instill an enthalpy change, Delta E 1 must be analyzed with steady flow equations and design tools at this preliminary level as are known in thermodynamics and propulsion dynamics and considering improvements in the power equation of the comoporession machine in question via evaluation of steady flow in arid out of a component compressor as shown in Figure 5.
  • the Euler turbine equation relates the power added to or removed from the flow, to characteristics of a rotating blade row.
  • the equation is based on the concepts of conservation of angular momentum and conservation of energy.
  • a representative model of the blade row describing representative vectors and metrics:
  • the Euler Turbomachinary Equation relates the temperature ratio (and hence the pressure ratio) across a compressor to the rotational speed and the change in momentum per unit mass.
  • the velocities used in this equation are what are denoted as absolute frame velocities (as opposed to relative frame velocities).
  • An axial compressor is typically made up of many alternating rows of rotating and stationary blades called rotors and stators, respectively, as shown.
  • the first stationary row (which comes in front of the rotor) is typically called the inlet guide vanes or IGV.
  • IGV inlet guide vanes
  • Each successive rotor-stator pair is called a compressor stage.
  • compressors with many blade rows are termed multistage compressors.
  • the rotor adds swirl to the flow, thus increasing the total energy carried in the flow by increasing the angular momentum (adding to the kinetic energy associated with the tangential or swirl velocity, 1/2H/ 2 ).
  • the stator removes swirl from the flow, but it is not a moving blade row and thus cannot add any net energy to the flow.
  • every stator row is a slower moving airfoil blade row, thus having the capacity to add net energy to the flow, as well as acting as a conversion device to the flow, adding some kinetic energy to the flow and raising the static pressure simultaneously of the flow.
  • Typical velocity and pressure profiles through a multistage axial compressor look like those shown in Figure 1.43.
  • a typical velocity and pressure profiles through an electric multistage ringmotor axial compressor are exhibited in Figure 1.44 where pressure rise is greater and velocity drop and Mach number are reduced across the compressor.
  • the IGV also adds no energy to the flow. It is designed to add swirl in the direction of rotor motion to lower the Mach number of the flow relative to the rotor blades, and thus improve the aerodynamic performance of the rotor.
  • Velocity triangles are typically used to relate the flow properties and blade design parameters in the relative frame (rotating with the moving blades), to the properties in the stationary or absolute frame.
  • Velocity triangles for an axial compressor stage Primed quantities are in the relative frame, un primed quantities are in the absolute frame.
  • the propulsive efficiency of a simple turbojet can be improved by extracting a portion of the energy from an engine's gas generator to drive a ducted propeller, called a fan.
  • the ducted propeller pushes a portion of the overall air through the turbine, but by-passes the turbine, exhausting to the rear at ambient air conditions.
  • the fan increases the propellant mass flow rate with an accompanying decrease in the required propellant exit velocity for a given thrust. Since the rate of production of "wasted" kinetic energy in the exit propellant gases varies as the first power with mass flow rate and as the square of the exit velocity, the net effect of increasing mass flow rate and decreasing the exit velocity is to reduce the wasted kinetic energy production and to improve the propulsive efficiency.
  • turbomachinary design has moved to supersonic low-bypass jet engine designs, whereby the bypass fan is reduced in size compared to a pure turbofan to maintain a relatively high mass air flow and exhaust velocity Mach number.
  • the approach offers greater efficiency through moderation of typically high endothermic and entropic thermal reactions of pure turbojets by optimizing mass flow rates and exhaust velocities.
  • turbofan stage(s) enables the turbine to be refined to the cruise flight condition and low-speed flight conditions by utilizing more of the combustion gases efficiently and by reducing the wasted kinetic energy. Improvements can be observed in a ring motor turbofan where it is not constrained by the available rotating speeds in a multi-shaft turbine design as it is rim driven and enables the fan stage to optimize and maximize typically constrained design variables as follows: optimized design in turbomachinary is focused on "ideal" mass flow through the engine core and the fan. In current turbofan designs, or supersonic low-bypass turbine designs the temperature drop through the turbine is greater than the temperature rise through the compressor since the turbine drives the fan in addition to the compressor.
  • the temperature drop across the fan can be minimized as compared to across the compressor, as this is beneficial in maintaining temperature during compression and assists in the entropic and endothermic reactions in the atomization of fuel in the combustor, subsequently mass flow of the fan can be increased relative to the compressor, more air can be compressed, Delta M over Delta C at any given T.
  • a drive shaft there remains a load on the turbine, in the form of a future design iteration for an electric generation source in the form of a turbine ring generator, which causes an electric load on the turbine machine invention.
  • Mass flow can be increased since stage loading for each compressor stage can be increased in an electric compressor as previously discussed (mass flow increases load capacity).
  • a ring motor electric bypass fan in the SonicBlue configuration of superconducting electromagnetics and magnetically levitated compressor offers zero electric resistance and zero drag. This further adds to the ability of the invention to mass load the turbo fan with inlet air beyond current design levels, thus increasing over all mass flow in the engine.
  • a turbofan engine is presented with an electric turbofan upstream of a multiaxial electric compressor as previously described in this paper.
  • the core flow and bypass flow are mixed together through an afterburner and nozzle.
  • Figure XXX a general form of relationship between flow area and Mach number of a Turbofan (does not account for stagnation condition at the IGV of the compressor).
  • turbofan cycle with mechanical compressor and mixed stream with afterburner is shown in the T- s Diagram.
  • Fluid dynamics requires equal static pressures at stations 6 and 16.
  • Normal design of the mixer has the mach numbers of the two entering streams equal.
  • the Mach numbers of the two respective streams can be matched, thus reducing boundary layer drag at the mixer wall, unsteady enthalpic mixing currents mid-stream, and the two pressures of the entering streams can be made equal, thus converse to mechanically driven designs total pressure ratio of the mixer can be brought to unity creating an ideal low- bypass turbofan engine with fan and compressor driven electrically.
  • the fan stream of the turbofan contains a fan rather than a compressor and does not have either a combustor or a turbine.
  • the turbofan sits upstream of the compressor, its ambient temperature flow mixing downstream outside of the combustor and ahead of the afterburner.
  • Current art in turbomachinary design of a mixed flow turbofan engine with afterburner as shown in Figure XXX using mechanical linkages (drive shaft) versus electrical load linkages as in the current invention prevent any management of gas mixing in the mixer area just described. Since velocity of bypass air and compressor air can be controlled electrically through RPM, the mixing process can be optimized. Further the management of the mixing process in this type of turbine proposed in the invention can have a positive effect on the combustion forming process adjacent to the mixer behind the turbine.
  • the invention described herein demonstrates that a multi-disc, turbofan assembly concept, because each fan disc is driven independently by an electric ring motor, the fan pressure ratio (hence the mass flow) and the bypass ratio can be varied and optimized against temperature across the main components, fan, compressor and turbine.
  • An integral expression of an "electric variable ratio bypass fan” with “bypass flow” in a mixed flow afterburning turbofan, as it relates to pressure and temperature, is described as: Tc - Tf

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

L'invention porte sur une soufflante de compresseur de turbine électrique pour propulsion hybride, le compresseur contenant un ou plusieurs étages rotors (compresseur et diffuseur), chacun d'eux étant entraîné par au moins un moteur électrique à bagues, de sorte que les étages rotors du compresseur sont conçus et adaptés de manière plus précise au rapport de compression à atteindre, dans les limites des caractéristiques de fonctionnement, des exigences de poussée et du domaine de vol du modèle de turbine.
EP07716374A 2006-01-09 2007-01-09 Soufflante de dérivation et compresseur de turbine électrique pour propulsion hybride Withdrawn EP1977082A2 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US75736906P 2006-01-09 2006-01-09
PCT/US2007/000307 WO2007081817A2 (fr) 2006-01-09 2007-01-09 Soufflante de dérivation et compresseur de turbine électrique pour propulsion hybride

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EP1977082A2 true EP1977082A2 (fr) 2008-10-08

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EP (1) EP1977082A2 (fr)
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WO (1) WO2007081817A2 (fr)

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WO2013102169A2 (fr) * 2011-12-30 2013-07-04 Rolls-Royce North American Technologies Inc. Soufflante à utiliser avec un séparateur air-particules
WO2013102113A2 (fr) 2011-12-30 2013-07-04 Rolls-Royce North American Technologies Inc. Moteurs à turbine à gaz comprenant des turbines à vitesse variable
CN113673060B (zh) * 2021-08-26 2024-02-27 上海交通大学 多级压气机模化方法及系统
CN114329771B (zh) * 2021-12-14 2024-02-23 西北工业大学 考虑叶片安装角的轴流压气机尺寸计算方法

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GB2356684A (en) * 1999-11-24 2001-05-30 Lorenzo Battisti Boundary layer control using electroformed microporous material

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See references of WO2007081817A2 *

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RU2008132808A (ru) 2010-02-20
WO2007081817A2 (fr) 2007-07-19
WO2007081817A3 (fr) 2007-12-06

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