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Preliminary assembly and initial test results of gas bearing gyroscope with high-g and sterilization capabilities
High gain, high frequency flexible responses in gimbaled multibody systems are discussed. Their origin and physical significance are described in terms of detailed mass and stiffness modeling at actuator/sensor interfaces. Guyan Reduction, Generalized Dynamic Reduction, inadequate mass modeling detail, as well as system mode truncation, are shown to suppress the high gain high frequency response and thereby lose system flexibility important for stability and performance predictions. Model validation by modal survey testing is shown to risk similar loss of accuracy. Difficulties caused by high frequency responses in component mode simulations, such as DISCOS, and also linearized system mode simulations, are described, and approaches for handling these difficulties are discussed.
To test liquid rocket engines (LREs) for combustion instabilities, devices such as bombs are often used to create pressure wave disturbances. Bombs, while effective, are inherently dangerous, expensive, and difficult to procure. Over the years, pulse guns have been used as a safer and more cost-effective way to generate controlled pressure disturbances in engine chambers. In anticipation of the need for stability testing of prototype LREs at NASA Marshall Space Flight Center (MSFC), a set of pulse guns have been designed, fabricated, tested, and characterized. The pulse gun program is funded by the RS-25 Engine Program managed out of MSFC and funded by NASA’s Space Launch System (SLS) through the MSFC Liquid Engines Office (LEO). A pulse gun is a simple device – like a traditional gun, it has a breech and barrel section. However, unlike a traditional gun, there is no bullet, as the purpose of the pulse gun is strictly to fire a high pressure pulse. Instead of a firing pin and primer that would normally be used in a traditional gun, an initiator is used to activate the gun powder. The initiator is a highly reliable pyrotechnic initiator. For this study, clone versions of the NASA Standard Initiator (NSI) were used. The initiator is used to ignite a pre-measured amount of gun powder loaded into the breech. When the pressure of the burnt gun powder exceeds the set point of a downstream burst disk, the disk ruptures, allowing the high pressure pulse to travel downstream through the barrel section. A ballistic pressure transducer located in the breech section is used to measure the short duration, high pressure pulse. Some configurations of the pulse gun have barrel sections that include one or more additional ballistic pressure transducers. These additional pressure measurements help track the degradation and damping of the pulse as it travels out of the barrel section. The measurements may also be used to determine the velocity of pulse. The objective of this paper is to present the different variants of this newly-developed pulse gun and characterize performance over a range of parameters. The parameters include breech diameter, barrel length, amount of gun powder used, the way the gun powder is packed, and the pressure setting of the burst disk. There are a total of six pulse guns: three with the 0.25 inch breech and three different length barrels, and three with the 0.40 inch breech and three different length barrels. For both breech sizes, the associated barrels are referred to as standard, one-port, and two-port barrels. The standard barrel has no instrumentation and is likely the barrel that will be used for engine stability testing. The one- and two-port barrels were designed specifically for pulse gun component testing to allow measurements of the magnitude and timing of the high pressure pulse as it makes its way through the pulse gun. The burst disks tested were commercially-available burst disks designed to rupture at 8,000, 16,000, and 24,000 psid. Testing was accomplished by firing the pulse gun into a test chamber pressurized with nitrogen at about 2300 psig. A total of four Model 113B23 High frequency ICP® pressure sensors (10k psi) were mounted in the test chamber, in the same plane as the pulse gun. Two of the sensors had “trimmed” adapters, and two did not. Given the symmetrical configuration of the sensors within the test chamber, different amplitudes of pressure measurements are attributed to the use of trimmed versus untrimmed adapters. The untrimmed adapters, with their narrower passages, tended to amplify the pressure amplitudes by as much as 50%. In total, 41 pulse gun tests have been conducted.Data are still being analyzed, but some trends are apparent. For example, measurements taken within the pulse gun are shown in Figures 3 and 4 for Hot-fires (HFs) # 11 and 39, respectively. Both tests were identical in that the 0.40 inch ID breech, two-port barrel, 8,000 psid burst disk, and same amount of gunpowder (6.639 grains for HF# 11, and 6.576 grains for HF# 39) were used, and the back pressure in the nitrogen test chamber was ~2300 psig. Despite keeping all these variables constant, the results from these two tests look quite different. For both tests, the first pressure peak shown in the P2111 trace within the breech is the firing of the initiator. The second peak and any subsequent peaks are from the combustion of the gunpowder. It should be noted that because the sensors are dynamic, the ~2300 psig baseline static pressure is not shown in the figures. For HF# 11, the burst disk clearly ruptured in a little over 0.1 ms from the time the initiator was fired. This is evident from the abrupt rise in pressure for P2112 followed by another abrupt rise in pressure for P2113. The sensors track the movement of the high pressure pulse through the pulse gun. For HF# 39, however, there was nearly 1 ms delay between the initiator firing and the burst disk rupturing. The only difference between these two tests was the way in which the gunpowder was packed within the breech. For HF# 11, the gunpowder was poured directly into the breech, atop the initiator, and held in place with a vegetable fiber wad appropriately sized for the barrel diameter. Once in place, the wad firmly held the gunpowder in place. For HF# 39, the gunpowder was rolled within cigarette paper with the ends of the paper twisted. The rolled gunpowder was gently pushed down into the breech, toward the initiator, and no wad was used. This method of loading the gunpowder consistently produced longer delays in the rupture of the burst disk, presumably because there was a steady, fuller burning of the gunpowder. The burst disk was rated for 8,000 psid, which means that the burst disk should not have opened until the pressure in the breech reached ~10,300 since there was ~2300 psig back pressure in the test chamber. In the case of HF# 11, the breech pressure only reached roughly 4200 psi before there was leakage either through or past the burst disk, but for HF#39, the breech pressure reached 10,600 psi prior to the burst disk opening. The temperature of the burst disk is certainly a factor in determining when the burst disk will rupture. A sharp rise in the burst disk temperature could weaken it and cause it to stray from the designed set point burst pressure. The burst disks are manufactured of Inconel to try to reduce their sensitivity to temperature, but how the burst disks perform above 900°F is not documented. All burst disks were engineered and tested by the manufacturer for ambient temperature conditions. One theory for why rolling the gunpowder in cigarette paper may produce better results is related to a more controlled temperature environment. The cigarette paper may act as a sheath that protects the burst disk from excessive temperatures during the rapid buildup of pressure in the breech. The ultimate benefit of not breeching the burst disk prematurely is shown in Figures 5 and 6, which graphs the resultant pressure pulse produced in the test chamber. For HF #11, the peak pressure on the P2123 trimmed adapter is about 270 psi, while for HF# 39 it is 343 psi, which is about 1.27 times greater in magnitude. In either case, the pressure pulse is ~20 μs in width once it reaches the test chamber. Another pressure rise about 100 μs later can be seen on the sensors located across from the pulse gun port, P2124 and P2121. These pressures are lower due to attenuation of the pressure wave as it moves across the test chamber. After the wave encounters the wall, it reflects back and forth within the chamber until it completely dissipates.
Testing for the low cycle fatigue and creep fatigue resistance of superalloys containing reduced amounts of cobalt is described. The test matrix employed involves a single high temperature appropriate for each alloy. A single total strain range, again appropriate to each alloy, is used in conducting strain controlled, low cycle, creep fatigue tests. The total strain range is based upon the level of straining that results in about 10,000 cycles to failure in a high frequency (0.5 Hz) continuous strain-cycling fatigue test. No creep is expected to occur in such a test. To bracket the influence of creep on the cyclic strain resistance, strain hold time tests with ore minute hold periods are introduced. One test per composition is conducted with the hold period in tension only, one in compression only, and one in both tension and compression. The test temperatures, alloys, and their cobalt compositions that are under study are given.
A ridge waveguide distributed feedback laser was developed in InGaAsP. These devices have demonstrated CW output powers over 7 mW with threshold currents as low as 60 mA at 25 C. Measurements of the frequency response of these devices show a 3 dB bandwidth of about 2 GHz, which may be limited by the mount. The best devices have a single mode spectra over the entire temperature range tested with a side mode suppression of about 20 dB in both CW and pulsed modes. The design of this device, including detailed modeling of the ridge guide structure, effective index calculations, and a discussion of the grating configuration are presented. Also, the fabrication of the devices is presented in some detail, especially the fabrication of and subsequent growth over the grating. In addition, a high frequency fiber pigtailed package was designed and tested, which is a suitable prototype for a commercial package.
The electrohydraulically actuated slotted-plate bypass valve used to control the position of the normal shock during wind-tunnel investigations of supersonic inlets also has proven to be a valuable device for determining inlet dynamics and creating airflow disturbances. Operation of previous valves at high frequencies (to 100 Hz) for extended testing has resulted in numerous failures. An improved bypass-valve design is presented which increases the cyclic tolerance of the device considerably over past designs. The use of dynamic limit criteria to obtain an optimum actuator-piston size results in a frequency response which is flat within + or - 3 decibels to 120 Hz for a peak-to-peak variation of 20 percent of full area.
A combination of analytic modeling and system identification methods have been used to develop an improved dynamic model describing the response of articulated rotor helicopters to control inputs. A high-order linearized model of coupled rotor/body dynamics including flap and lag degrees of freedom and inflow dynamics with literal coefficients is compared to flight test data from single rotor helicopters in the near hover trim condition. The identification problem was formulated using the maximum likelihood function in the time domain. The dynamic model with literal coefficients was used to generate the model states, and the model was parametrized in terms of physical constants of the aircraft rather than the stability derivatives resulting in a significant reduction in the number of quantities to be identified. The likelihood function was optimized using the genetic algorithm approach. This method proved highly effective in producing an estimated model from flight test data which included coupled fuselage/rotor dynamics. Using this approach it has been shown that blade flexibility is a significant contributing factor to the discrepancies between theory and experiment shown in previous studies. Addition of flexible modes, properly incorporating the constraint due to the lag dampers, results in excellent agreement between flight test and theory, especially in the high frequency range.
A combination of analytic modeling and system identification methods have been used to develop an improved dynamic model describing the response of articulated rotor helicopters to control inputs. A high-order linearized model of coupled rotor/body dynamics including flap and lag degrees of freedom and inflow dynamics with literal coefficients is compared to flight test data from single rotor helicopters in the near hover trim condition. The identification problem was formulated using the maximum likelihood function in the time domain. The dynamic model with literal coefficients was used to generate the model states, and the model was parametrized in terms of physical constants of the aircraft rather than the stability derivatives, resulting in a significant reduction in the number of quantities to be identified. The likelihood function was optimized using the genetic algorithm approach. This method proved highly effective in producing an estimated model from flight test data which included coupled fuselage/rotor dynamics. Using this approach it has been shown that blade flexibility is a significant contributing factor to the discrepancies between theory and experiment shown in previous studies. Addition of flexible modes, properly incorporating the constraint due to the lag dampers, results in excellent agreement between flight test and theory, especially in the high frequency range.
This paper presents a novel means for suppressing gear mesh-related vibrations. The key components in this approach are piezoelectric actuators and a high-frequency, analog feed-forward controller. Test results are presented and show up to a 70-percent reduction in gear mesh acceleration and vibration control up to 4500 Hz. The principle of the approach is explained by an analysis of a harmonically excited, general linear vibratory system.
Development of a rover deployable Ground Penetrating Radar (GPR) involves: the nearly finished design and testing of a transducer array with high frequency (bistatic) and low frequency (monostatic) components; and design and development of a complete impulse GPR system.
Computer Generated Holograms (CGHs) have been used for some time to calibrate interferometers that require nulling optics. A typical scenario is the testing of aspheric surfaces with an interferometer placed near the paraxial center of curvature. Existing CGH technology suffers from a reduced capacity to calibrate middle and high spatial frequencies. The root cause of this shortcoming is as follows: the CGH is not placed at an image conjugate of the asphere due to limitations imposed by the geometry of the test and the allowable size of the CGH. This innovation provides a calibration system where the imaging properties in calibration can be made comparable to the test configuration. Thus, if the test is designed to have good imaging properties, then middle and high spatial frequency errors in the test system can be well calibrated. The improved imaging properties are provided by a rudimentary auxiliary optic as part of the calibration system. The auxiliary optic is simple to characterize and align to the CGH. Use of the auxiliary optic also reduces the size of the CGH required for calibration and the density of the lines required for the CGH. The resulting CGH is less expensive than the existing technology and has reduced write error and alignment error sensitivities. This CGH system is suitable for any kind of calibration using an interferometer when high spatial resolution is required. It is especially well suited for tests that include segmented optical components or large apertures.
The acoustic consequences of sealing the Helmholtz resonators of the NASA Lewis 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT) were experimentally evaluated. This resonator sealing was proposed in order to avoid entrapment of hydrogen during tests of advanced hydrogen-fueled engines. The resonators were designed to absorb energy in the 4- to 20-Hz range; thus, this investigation is primarily concerned with infrasound. Limited internal and external noise measurements were made at tunnel Mach numbers ranging from 0.5 to 2.0. Although the resonators were part of the acoustic treatment installed because of a community noise problem their sealing did not seem to indicate a reoccurrence of the problem would result. Two factors were key to this conclusion: (1) A large bulk treatment muffler downstream of the resonators was able to make up for much of the attenuation originally provided by the resonators, and (2) there was no noise source in the tunnel test section. The previous community noise problem occurred when a large ramjet was tested in an open-loop tunnel configuration. If a propulsion system producing high noise levels at frequencies of less than 10 Hz were tested, the conclusion on community noise would have to be reevaluated.
The development of advanced high performance constant-volume-combustion-cycle engines (CVCCE) requires robust design of the engine components that are capable of enduring harsh combustion environments under high frequency thermal and mechanical fatigue conditions. In this study, a simulated engine test rig has been established to evaluate thermal fatigue behavior of a candidate engine combustor material, Haynes 188, under superimposed CO2 laser surface impulsive thermal loads (30 to 100 Hz) in conjunction with the mechanical fatigue loads (10 Hz). The mechanical high cycle fatigue (HCF) testing of some laser pre-exposed specimens has also been conducted under a frequency of 100 Hz to determine the laser surface damage effect. The test results have indicated that material surface oxidation and creep-enhanced fatigue is an important mechanism for the surface crack initiation and propagation under the simulated CVCCE engine conditions.
Gimbaled spinmotor, gyro assembly, and gyro test results for gas bearing gyroscopes with high G and sterilization capability
NASA is pursuing the development of electrified propulsion (EP) technologies to improve air transportation in terms of efficiency, affordability, and sustainability. The power requirements are expected to reach 20 megawatts for large EP commercial airliners. A key challenge is transmission and distribution of the high voltage (20 kilovolts) needed for some EP systems. This is a challenge because the risk of electrical failure in aircraft at high altitude (low pressure) due to corona discharge and other forms of partial discharge increases significantly when voltages exceed 327 V. Corona discharge is a form of partial discharge in which gaseous molecules are ionized by strong electric fields. Corona discharge leads to aging in power transmission lines. Aging is a leading cause of electrical failure in electrical insulation materials. In particular, aging from electrical, vibrational, and thermal stresses decrease the performance life of insulation materials. There are currently no test standards or equipment that can effectively age materials under high altitude, voltage and frequency conditions. This poster proposes a design for a test system including an environmental chamber that can simulate the environment of future EP systems for material and component aging tests.
NASA's Marshall Space Flight Center(MSFC) has been tasked with developing a 60,000 pound thrust, pump-fed, LOX/RP-1 engine under the Advanced Space Transportation Program(ASTP). This government-led design has been designated the Fastrac engine. The X-34 vehicle will use the Fastrac engine as the main propulsion system. The X-34 will be a suborbital vehicle developed by the Orbital Sciences Corporation. The X-34 vehicle will be launched from an L-1011 airliner. After launch, the X-34 vehicle will be able to climb to altitudes up to 250,000 feet and reach speeds up to Mach 8, over a mission range of 500 miles. The overall length, wingspan, and gross takeoff weight of the X-34 vehicle are 58.3 feet, 27.7 feet and 45,000 pounds, respectively. This report summarizes the plan of achieving a Fastrac thrust chamber assembly(TCA) stable bomb test that meets the JANNAF standards, the Fastrac TCA design, and the combustion instabilities exhibited by the Fastrac TCA during testing at MSFC's test stand 116 as determined from high-frequency fluctuating pressure measurements. This report also summarizes the characterization of the combustion instabilities from the pressure measurements and the steps taken to eliminate the instabilities.
The geometric characteristics of LANDSAT 4 images were studied. The extent of image distortion caused by the physical process of writing the LANDSAT 4 images on film was determined. The geometric accuracies inherent in the digital images themselves were characterized. Test materials consisted of film images of test targets generated by the laser beam recorders (LBR), the Optronics Photowrite film writer, and digital image films of a strip 600 lines deep across the full width of band 5 of the Washington, D.C. Thematic Mapper scene. The tests were made by least squares adjustment of an array of measured image points to a corresponding array of control points. The film test array consists of 25 test points, and the image test array consists of 60 test points, resulting in a model that can be used to characterize the low frequency, random geometric distortions of the images. The tests were not suitable for examining jitter or similar high-frequency distortions.
An overview is presented of some of the hypersonic technology that will become the baseline for more advanced commercial aerospace systems and new military transportation systems for carrying astronauts and equipment into space. Attention is given to the X-15 aeronautical research program, the X-20 DYNA-SOAR, and the current X-30 National Aerospace Plane. Consideration is given to FEM analysis methods, modal testing conducted to measure the structure's resonant frequencies, dampings, and mode shapes, and high-temperature, high-speed wind tunnel testing and in-flight measurement of steady and unsteady pressures at Mach 3 and above.