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Proposed tachometer produces voltages proportional to shaft angular velocity and (by differentiation) acceleration. Coil moving in homopolar field generates emf proportional to shaft angular velocity.
Methods for a blood pump bearing system within a pump housing to support long-term high-speed rotation of a rotor with an impeller blade having a plurality of individual magnets disposed thereon to provide a small radial air gap between the magnets and a stator of less than 0.025 inches. The bearing system may be mounted within a flow straightener, diffuser, or other pump element to support the shaft of a pump rotor. The bearing system includes a zirconia shaft having a radiused end. The radiused end has a first radius selected to be about three times greater than the radius of the zirconia shaft. The radiused end of the zirconia shaft engages a flat sapphire endstone. Due to the relative hardness of these materials a flat is quickly produced during break-in on the zirconia radiused end of precisely the size necessary to support thrust loads whereupon wear substantially ceases. Due to the selection of the first radius, the change in shaft end-play during pump break-in is limited to a total desired end-play of less than about 0.010 inches. Radial loads are supported by an olive hole ring jewel that makes near line contact around the circumference of the shaft to support high speed rotation with little friction. The width of olive hole ring jewel is small to allow heat to conduct through to thereby prevent heat build-up in the bearing. A void defined by the bearing elements may fill with blood that then coagulates within the void. The coagulated blood is then conformed to the shape of the bearing surfaces.
The vertical take-off and landing (VTOL) and high-speed cruise capability of the NASA Large Civil Tilt-Rotor (LCTR) notional vehicle is envisaged to enable increased throughput in the national airspace. A key challenge of the LCTR is the requirement to vary the main rotor speeds from 100% at take-off to near 50% at cruise as required to minimize mission fuel burn. The variable-speed power-turbine (VSPT), driving a fixed gear-ratio transmission, provides one approach for effecting this wide speed variation. The key aerodynamic and rotordynamic challenges of the VSPT were described in the FAP Conference presentation. The challenges include maintaining high turbine efficiency at high work factor, wide (60 deg.) of incidence variation in all blade rows due to the speed variation, and operation at low Reynolds numbers (with transitional flow). The PT -shaft of the VSPT must be designed for safe operation in the wide speed range required, and therefore poses challenges associated with rotordynamics. The technical challenges drive research activities underway at NASA. An overview of the NASA SRW VSPT research activities was provided. These activities included conceptual and preliminary aero and mechanical (rotordynamics) design of the VSPT for the LCTR application, experimental and computational research supporting the development of incidence tolerant blading, and steps toward component-level testing of a variable-speed power-turbine of relevance to the LCTR application.
The flexural vibrations of a rotating shaft, running through one or more critical speeds, can be reduced to an acceptably low level by applying suitable control forces at an intermediate span position. If electromagnets are used to produce the control forces then it is possible to implement a wide variety of control strategies. A test rig is described which includes a microprocessor-based controller, in which such strategies can be realised in terms of software-based algorithms. The electromagnet configuration and the method of stabilising the electromagnet force-gap characteristic are discussed. The bounds on the performance of the system are defined. A simple control algorithm is outlined, where the control forces are proportional to the measured displacement and velocity at a single point on the shaft span; in this case the electromagnet behaves in a similar manner to that of a parallel combination of a linear spring and damper. Experimental and predicted performance of the system are compared, for this type of control, where various programmable rates of damping are applied.
The Space Shuttle body flap (BF) is located beneath the main engine nozzles and is required for proper aerodynamic control during orbital descent. The body flap is controlled by four actuators connected by a common shaft and driven by the hydraulic power drive unit. Inspection of the actuators during refurbishment revealed three shaft bearings with unexpected damage. One was coated with black oxide on the balls and race wear surfaces, a second contained a relatively deep wear scar, and the third with scratches and an aluminum particle in the wear track. A shaft bearing life test program was initiated to measure the wear life and explain the 5.08-micrometer wear scar. A tribological analysis was conducted to demonstrate that the black oxide coated wear surfaces did not damage the bearing, interfere with the lubrication, or cause severe bearing wear. Pre-damaged (equivalent of 30 missions), commercial equivalent bearings and previously flown shaft bearings were tested at axial loads, speeds, and temperatures seen during flight operations. These bearing were successfully life tested at 60 C for 24 hours or 90 flights. With a safety factor of 4X, the bearings were qualified for 22 flights when only a maximum of 12 flights are expected. Additional testing at 23 C was performed to determine the lubricant life and to further understand the mechanism that caused the blackened balls. Test results indicating bearing life was shortened at a lower temperature surprised the investigators. Start\Stop bearing testing that closely simulates mission profile was conducted at 23 C. Results of this testing showed lubricant life of 12 flights including a safety factor of four. Additional testing with bearings that have the equivalent of 30 missions of damage is being tested at 23 C. These tests are being performed over the Shuttle load profile to demonstrate the residual bearing life in the actuators exceeds 12 missions. Testing showed that the end of the shaft bearing life was characterized by bearing temperature rise, preload drop, and the onset of a severe wear bearing failure mechanism. The severe wear failure mechanism is characterized by rough wear scars, extensive bearing wear and steel transfer between the balls and the races.
The Brayton rotating unit (BRU), operating on a gas bearing system, has been designed, fabricated, and demonstrated for use in a closed Brayton cycle space power conversion system. The BRU uses a binary mixture of xenon and helium (molecular weight, 83.8) as the cycle working fluid and bearing lubricating medium and was designed to produce from 2.25 to 10.5 kw sub e of 1200 Hz three-phase electrical power. The single-shaft rotating assembly operates at a design speed of 36,000 rpm and comprises a radial single-stage compressor, a four-pole Rice alternator rotor, and a radial inflow turbine. Four units, a dynamic simulator and three component research packages, were supplied to the NASA Lewis Research Center for performance testing and further development.
An unstable asynchronous vibration problem in a high pressure centrifugal compressor and the remedial actions against it are described. Asynchronous vibration of the compressor took place when the discharge pressure (Pd) was increased, after the rotor was already at full speed. The typical spectral data of the shaft vibration indicate that as the pressure Pd increases, pre-unstable vibration appears and becomes larger, and large unstable asynchronous vibration occurs suddenly (Pd = 5.49MPa). A computer program was used which calculated the logarithmic decrement and the damped natural frequency of the rotor bearing systems. The analysis of the log-decrement is concluded to be effective in preventing unstable vibration in both the design stage and remedial actions.
Convertible engine aircraft gas-turbine is two engines in one. Produces turbofan thrust, turboshaft power, or any combined thrust and shaftpower continously while operating up to full speed. Used to power vertical/short-takeoff-and-landing (V/STOL) airplanes and advanced high-speed rotorcraft. Rotorcraft shows use of convertible engines rather than separate engines for rotor power and forward thrust affords advantages in installation and save 16 percent in fuel and 20 percent in direct operating costs. Engine used for propulsion of new high-speed rotorcraft that needs both thrust and shaft power. Also used to cross-couple fans of two-engine V/STOL aircraft.
The torque-limited touchdown bearing system (TLTBS) is a backup mechanical-bearing system for a high-speed rotary machine in which the rotor shaft is supported by magnetic bearings in steady-state normal operation. The TLTBS provides ball-bearing support to augment or supplant the magnetic bearings during startup, shutdown, or failure of the magnetic bearings. The TLTBS also provides support in the presence of conditions (in particular, rotational acceleration) that make it difficult or impossible to control the magnetic bearings or in which the magnetic bearings are not strong enough (e.g., when the side load against the rotor exceeds the available lateral magnetic force).
Far-field acoustic power level and performance analyses of open rotor model F31/A31 have been performed to determine its noise characteristics at simulated scaled takeoff, nominal takeoff, and approach flight conditions. The nonproprietary parts of the data obtained from experiments in 9- by 15-Foot Low-Speed Wind Tunnel (915 LSWT) tests were provided by NASA Glenn Research Center to perform the analyses. The tone and broadband noise components have been separated from raw test data by using a new data analysis tool. Results in terms of sound pressure levels, acoustic power levels, and their variations with rotor speed, angle of attack, thrust, and input shaft power have been presented and discussed. The effect of an upstream pylon on the noise levels of the model has been addressed. Empirical equations relating model's acoustic power level, thrust, and input shaft power have been developed. The far-field acoustic efficiency of the model is also determined for various simulated flight conditions. It is intended that the results presented in this work will serve as a database for comparison and improvement of other open rotor blade designs and also for validating open rotor noise prediction codes.
NASA is supporting the development of Electrified Aircraft Propulsion (EAP) technology due to its potential to reduce aircraft fuel burn, emissions, and noise as well as improving safety and performance. One focus of this research is the electrification of conventional turbomachinery propulsion systems, which offers ways to improve the performance and operability of turbine-engine powered aircraft through the addition of electro-mechanical systems. These hybrid-electric turbine engines provide additional actuation and energy management control opportunities for improving stability and transient response behavior. This paper summarizes the results of a Hardware-in-the-Loop (HIL) test performed at the NASA Electric Aircraft Testbed (NEAT) during the summer of 2022. The test demonstrates the feasibility and performance of an advanced energy management control strategy by integrating a simulated turbofan engine with scaled electro-mechanical hardware. A full-scale real-time reference model of a geared turbofan was run alongside a scaled electro-mechanical system representing the electrified turbofan components operating at a megawatt-scale power level. The model was interfaced with the hardware through a novel closed-loop control and scaling algorithm that emulated the dynamic speed and torque response of the turbofan shafts. The control strategy was implemented on the electrical machines connected to the emulated turbomachinery shafts. The results from the testbed are compared against simulations that predict the testbed and geared turbofan model operation. The energy management control strategy successfully changed the operating point of the engine model and improved its stability during throttle transients. These results also demonstrate the success of the novel closed loop control and scaling approach for emulating turbomachinery and elevate the Technology Readiness Level (TRL) of the energy management control strategy.
NASA is supporting the development of Electrified Aircraft Propulsion (EAP) technology due to its potential to reduce aircraft fuel burn, emissions, and noise as well as improving safety and performance. One focus of this research is the electrification of conventional turbomachinery propulsion systems, which offers ways to improve the performance and operability of turbine-engine powered aircraft through the addition of electro-mechanical systems. These hybrid-electric turbine engines provide additional actuation and energy management control opportunities for improving stability and transient response behavior. This paper summarizes the results of a Hardware-in-the-Loop (HIL) test performed at the NASA Electric Aircraft Testbed (NEAT) during the summer of 2022. The test demonstrates the feasibility and performance of an advanced energy management control strategy by integrating a simulated turbofan engine with scaled electro-mechanical hardware. A full-scale real-time reference model of a geared turbofan was run alongside a scaled electro-mechanical system representing the electrified turbofan components operating at a megawatt-scale power level. The model was interfaced with the hardware through a novel closed-loop control and scaling algorithm that emulated the dynamic speed and torque response of the turbofan shafts. The control strategy was implemented on the electrical machines connected to the emulated turbomachinery shafts. The results from the testbed are compared against simulations that predict the testbed and geared turbofan model operation. The energy management control strategy successfully changed the operating point of the engine model and improved its stability during throttle transients. These results also demonstrate the success of the novel closed loop control and scaling approach for emulating turbomachinery and elevate the Technology Readiness Level (TRL) of the energy management control strategy.
A 1-MW wellhead generator was tested in 1980, 1981, and 1982 by Mexico, Italy, and New Zealand at Cerro Prieto, Cesano, and Broadlands, respectively. The total flow helical screw expander portable power plant, Model 76-1, had been built for the U.S. Government and field-tested in Utah, USA, in 1978 and 1979. The expander had oversized internal clearances designed for self-cleaning operation on fluids that deposit adherent scale normally detrimental to the utiliation of liquid dominated fields. Conditions with which the expander was tested included inlet pressures of 64 to 220 psia, inlet qualities of 0% to 100%, exhaust pressures of 3.1 to 40 psia, electrial loads of idle and 110 to 933 kW, electrical frequencies of 50 and 60 Hz, male rotor speeds of 2500 to 4000 rpm, and fluid characteristics to 310,000 ppm total dissolved solids and noncondensables to 38 wt % of the vapor. Some testing was done on-grid. Typical expander isentropic efficiency was 40% to 50% with the clearances not closed, and 5 percentage points or more higher with the clearances partly closed. The expander efficiency increased approximately logarithmically with shaft power for most operations, while inlet quality, speed, and pressure ratio across the machine had only small effects. These findings are all in agreement with the Utah test results.
For a flexible shaft designed to pass through a number of bending critical speeds, a squeeze-film damper has been designed and tested. The damper properties were selected to provide control of all critical speeds, while meeting additional constraints of high power transmission requirements and damper simplicity. The damper was fabricated and installed and its ability to control flexible shaft vibrations was demonstrated by the comparison of vibration amplitudes both with and without the damper.
This paper focuses on the conceptual design of a two spool compressor for the NASA Large Civil Tilt Rotor engine, which has a design-point pressure ratio goal of 30:1 and an inlet weight flow of 30.0 lbm/sec. The compressor notional design requirements of pressure ratio and low-pressure compressor (LPC) and high pressure ratio compressor (HPC) work split were based on a previous engine system study to meet the mission requirements of the NASA Subsonic Rotary Wing Projects Large Civil Tilt Rotor vehicle concept. Three mean line compressor design and flow analysis codes were utilized for the conceptual design of a two-spool compressor configuration. This study assesses the technical challenges of design for various compressor configuration options to meet the given engine cycle results. In the process of sizing, the technical challenges of the compressor became apparent as the aerodynamics were taken into consideration. Mechanical constraints were considered in the study such as maximum rotor tip speeds and conceptual sizing of rotor disks and shafts. The rotor clearance-to-span ratio in the last stage of the LPC is 1.5% and in the last stage of the HPC is 2.8%. Four different configurations to meet the HPC requirements were studied, ranging from a single stage centrifugal, two axi-centrifugals, and all axial stages. Challenges of the HPC design include the high temperature (1,560deg R) at the exit which could limit the maximum allowable peripheral tip speed for centrifugals, and is dependent on material selection. The mean line design also resulted in the definition of the flow path geometry of the axial and centrifugal compressor stages, rotor and stator vane angles, velocity components, and flow conditions at the leading and trailing edges of each blade row at the hub, mean and tip. A mean line compressor analysis code was used to estimate the compressor performance maps at off-design speeds and to determine the required variable geometry reset schedules of the inlet guide vane and variable stators that would result in the transonic stages being aerodynamically matched with high efficiency and acceptable stall margins based on user specified maximum levels of rotor diffusion factor and relative velocity ratio.
A finite element model of a rotor bearing system was analyzed to determine the stability limits of the forward, backward, and centered Euler; Runge-Kutta; Milne; and Adams numerical integration techniques. The analysis concludes that the highest frequency mode determines the maximum time step for a stable solution. Thus, the number of mass elements should be minimized. Increasing the damping can sometimes cause numerical instability. For a uniform shaft, with 10 mass elements, operating at approximately the first critical speed, the maximum time step for the Runge-Kutta, Milne, and Adams methods is that which corresponds to approximately 1 degree of shaft movement. This is independent of rotor dimensions.
NASA and Pratt & Whitney will collaborate under a Space Act Agreement to perform spin testing of the impact damper to verify damping effectiveness and durability. Pratt & Whitney will provide the turbine blade and damper hardware for the tests. NASA will provide the facility and perform the tests. Effectiveness and durability will be investigated during and after sustained sweeps of rotor speed through resonance. Tests of a platform wedge damper are also planned to compare its effectiveness with that of the impact damper. Results from baseline tests without dampers will be used to measure damping effectiveness. The self-tuning impact damper combines two damping methods-the tuned mass damper and the impact damper. It consists of a ball located within a cavity in the blade. This ball rolls back and forth on a spherical trough under centrifugal load (tuned mass damper) and can strike the walls of the cavity (impact damper). The ball s rolling natural frequency is proportional to the rotor speed and can be designed to follow an engine-order line (integer multiple of rotor speed). Aerodynamic forcing frequencies typically follow these engineorder lines, and a damper tuned to the engine order will most effectively reduce blade vibrations when the resonant frequency equals the engine-order forcing frequency. This damper has been tested in flat plates and turbine blades in the Dynamic Spin Facility. During testing, a pair of plates or blades rotates in vacuum. Excitation is provided by one of three methods--eddy-current engine-order excitation (ECE), electromechanical shakers, and magnetic bearing excitation. The eddy-current system consists of magnets located circumferentially around the rotor. As a blade passes a magnet, a force is imparted on the blade. The number of magnets used can be varied to change the desired engine order of the excitation. The magnets are remotely raised or lowered to change the magnitude of the force on the blades. The other two methods apply force to the rotating shaft itself at frequencies independent of the rotor speed. During testing, blade vibration is monitored with strain gauges and laser displacement probes.
A conceptual design study was completed for a 360 kW Helium-Xenon closed Brayton cycle turbogenerator. The selected configuration is comprised of a single-shaft gas turbine engine coupled directly to a high-speed generator. The engine turbomachinery includes a 2.5:1 pressure ratio compression system with an inlet corrected flow of 0.44 kg/sec. The single centrifugal stage impeller discharges into a scroll via a vaned diffuser. The scroll routes the air into the cold side sector of the recuperator. The hot gas exits a nuclear reactor radiator at 1300 K and enters the turbine via a single-vaned scroll. The hot gases are expanded through the turbine and then diffused before entering the hot side sector of the recuperator. The single shaft design is supported by air bearings. The high efficiency shaft mounted permanent magnet generator produces an output of 370 kW at a speed of 60,000 rpm. The total weight of the turbogenerator is estimated to be only 123 kg (less than 5% of the total power plant) and has a volume of approximately 0.11 cubic meters. This turbogenerator is a key element in achieving the 40 to 45% overall power plant thermal efficiency.