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A Nd:YAG laser was end pumped with 2.2 W of continuous-wave (CW) diode laser output. Efficient operation of the laser at high repetition rates was emphasized. This laser provides 890 mW of TEM00 CW output at 1064 nm, and 340 mW of 532 nm average power at a Q-switched repetition rate of 25 kHz. Experimental data are compared with analysis.
The potential application of rotating detonation engines (RDEs) in rocket combustors hinges on a fundamental understanding of detonation wave structure and injector characteristics with fuel and oxidizer compositions relevant in rocket systems. Simultaneous 300 kHz-rate broadband OH* chemiluminescence and OH-PLIF imaging is employed in a fully optically accessible Natural Gas-GOx rotating detonation rocket engine (RDRE) to visualize reactant refill dynamics and detonation wave structure. A custom-built KTP-type optical parametric oscillator (OPO) is coupled with a nanosecond high-repetition-rate burst-mode laser to output284 nm light and target excitation of the Q1(9) transition in the OH radical. Significant deflagrative burning is observed throughout the chamber as a consequence of the oxygen-rich environment. Trailing Azimuthal Reflected Shock Combustion (ARSC) system, similar to those in a H2-air RDE are observed, burning unburned reactants in the region immediately following the primary detonation wave. Contact burning, as indicated in this study, does not seem to be a primary loss mechanism. The simultaneous measurement of OH and OH* show that axial locations exist in the refill process where OH radicals are present, and produced due to shear layer induced deflagration, however, these zones do not produce excited state OH*. While a deeper understanding of the underlying physics in RDRE systems requires further investigation, this work highlights a first-of-its-kind visualization of the turbulent combustion product field and reactant refill characteristics in this highly unsteady environment.
Selective two-photon absorptive resonance femtosecond laser electronic excitation tagging (STARFLEET) velocimetry is demonstrated for the first time in a NASA Langley wind tunnel with high repetition-rate and single-shot imaging. Experiments performed in the 0.3-meter Transonic Cryogenic Tunnel (TCT) allowed for testing at 300 K over a range of pressures (124 to 517 kPa) and Mach numbers (0.2-0.8) for freestream conditions and flow behind a cylindrical model. Measurement precision and accuracy are determined for the current set of experiments, as are signal intensity and lifetime. Precisions of 3-5 m/s (based on one standard deviation) were typical in the experiment; precisions better than 2% of the mean velocity were obtained for some of the highest velocity conditions. Agreement within a mean error of 3 m/s between STARFLEET freestream velocity measurements and facility DAS readings is demonstrated. STARFLEET is also shown to return spatially-resolved velocity profiles, though some binning of the signal is required.
The inductive pulsed plasma thruster (IPPT) is an electromagnetic plasma accelerator that has been identified in NASA roadmaps as an enabling propulsion technology for some niche low-power missions and for high-power in-space propulsion needs. The IPPT is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged producing a high current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. Thrusters of this type possess many demonstrated and potential benefits that make them worthy of continued investigation. The electrodeless nature of these thrusters eliminates the lifetime and contamination issues associated with electrode erosion in conventional electric thrusters. Also, a wider variety of propellants are accessible when compatibility with metallic electrodes in no longer an issue. IPPTs have been successfully operated using propellants like ammonia, hydrazine, and CO2, and there is no fundamental reason why they would not operate on other in situ propellants like H2O. It is well-known that pulsed accelerators can maintain constant specific impulse (I(sub sp)) and thrust efficiency (eta(sub t)) over a wide range of input power levels by adjusting the pulse rate to hold the discharge energy per pulse constant. It has also been demonstrated that an inductive pulsed plasma thruster can operate in a regime where eta(sub t) is relatively constant over a wide range of I(sub sp) values (3000-8000 s). Finally, thrusters in this class have operated in single-pulse mode at high energy per pulse, and by increasing the pulse rate they offer the potential to process very high levels of power using a single thruster. There has been significant previous research on IPPTs designed around a planar-coil (flat-plate) geometry. The most notable of these was the Pulsed Inductive Thruster (PIT), with the PIT MkV presently representing the state-of- the-art in pulsed high-power IPPT technological development. In this paper, we focus on two planar-geometry devices that operate at significantly different power levels. Most work performed at NASA-Marshall Space Flight Center (MSFC) has, to date, focused on lower power thruster operation (approx. = 10s to 100s of J/pulse, up to 2-2.5 kW average power throughput) and previously described. The most recent work aimed to assemble a device that could be tested in cyclic mode on a thrust-stand, and which could augment the existing data set for IPPTs. In addition, the thruster was designed to serve as a test-bed for solid state switching circuitry and pulsed gas valves, with the modular design of the device allowing for variation in or upgrades to test configuration. Recently, MSFC obtained on loan from the Georgia Institute of Technology (Atlanta, GA) the PIT MkVI, successor to the PIT MkV. The MkV and MkVI are similar in design with much of the hardware from the former, specifically the capacitors and spark-gap switches, being reused in the latter. The coil is similar in geometry but has bent copper rods used in the latest iteration in place of the Litz wire windings found in the MkV. The MkVI master switch for the spark gaps is located in the vacuum chamber contained within a sealed, pressurized vessel fastened to the back of the thruster. This is different from the MkV where many capacitor charging lines and spark gap-triggering delay lines ran to the thruster from a master trigger located outside the vacuum chamber. The MkVI was damaged during testing soon after its fabrication was completed. The thruster arrived at MSFC still-damaged and mostly disassembled into many individual pieces. The device has been repaired, with a few additional design changes implemented after discussions with the late Prof. Lovberg regarding the initial testing results and issues encountered. In the present work, we present results from testing of both the small IPPT and the larger MkVI thruster. The smaller device (Fig. 1) is tested on a thrust stand on multiple gases to demonstrate its capability to operate in a repetition-rate mode and serve as a IPPT technology-development testbed. The larger MkVI (Fig. 2) is operated for the first time in its newly reconstituted state, demonstrating full-power pulsed operation and, for the first time, repetition-rate operation of a high-power IPPT. The additional upgrades required for synchronous operation of all the pulsed systems in single-pulse and repetition-rate mode are described in detail.
In the present work, we use a burst-mode laser and optical parametric oscillator system to perform high-spectral resolution NO-PLIF measurements of an underexpanded jet at a repetition-rate of 100 kHz, with the motivation of multi-parameter measurements of temperature, pressure, and velocity. The laser frequency of the 1064 nm seed laser for the burst-mode laser was scanned during the burst to cover two neighboring absorption line pairs near 226 nm. The peak PLIF signal intensity varies along the axial (z) direction of the underexpanded jet as the laser frequency is scanned, which we attribute to the collisional shift induced by the flow and only revealed due to the narrow linewidth of the laser. A pseudo-Voight fit is applied to the LIF excitation spectra on a pixel-by-pixel basis to measure the spectral position of the peak intensity for the two transition pairs and their amplitude. The spectral position of the peak intensity is used to derive a frequency shift, which is separated into its collisional and Doppler components using the axisymmetric nature of the flow field. The amplitude of the measured peaks is used for two-line rotational thermometry. Challenges for quantitative measurements using such an approach are discussed, including measuring the spatial variations in the energy distribution of the laser sheet at a 100 kHz repetition rate and uncertainty/variability in the step size during the fast frequency scan.
In the present work, we use a burst-mode laser and optical parametric oscillator system to perform high-spectral resolution NO-PLIF measurements of an underexpanded jet at a repetition-rate of 100 kHz, with the motivation of multi-parameter measurements of temperature, pressure, and velocity. The laser frequency of the 1064 nm seed laser for the burst-mode laser was scanned during the burst to cover two neighboring absorption line pairs near 226 nm. The peak PLIF signal intensity varies along the axial (z) direction of the underexpanded jet as the laser frequency is scanned, which we attribute to the collisional shift induced by the flow and only revealed due to the narrow linewidth of the laser. A pseudo-Voight fit is applied to the LIF excitation spectra on a pixel-by-pixel basis to measure the spectral position of the peak intensity for the two transition pairs and their amplitude. The spectral position of the peak intensity is used to derive a frequency shift, which is separated into its collisional and Doppler components using the axisymmetric nature of the flow field. The amplitude of the measured peaks is used for two-line rotational thermometry. Challenges for quantitative measurements using such an approach are discussed, including measuring the spatial variations in the energy distribution of the laser sheet at a 100 kHz repetition rate and uncertainty/variability in the step size during the fast frequency scan.
Amplifier and stretcher system, which minimizes thermal effects and compensates for repetition-rate effects, maintains resolution levels in spectrum analysis. An additional inverting amplifier is used in the system to provide a noiseless charge restorer.
THE inductive pulsed plasma thruster (IPPT) is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged producing a high current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. In the present work, we present a summary of the IPPT research and development conducted at NASA's Marshall Space Flight Center (MSFC). As a higher-power, still relatively low readiness level system, there are many issues associated with the eventual deployment and use of the IPPT as a primary propulsion system on spacecraft that remain to be addressed. The present program aimed to fabricate and test hardware to explore how these issues could be addressed. The following specific areas were addressed within the program and will be discussed within this paper. a) Conical theta-pinch IPPT geometry thruster configuration. b) Repetition-rate multi-kW thruster pulsing. c) Long-lifetime pulsed gas valve. d) Fast pulsed gas valve driver and controller. e) High-voltage, repetitive capacitor charging power processing unit. During the course of testing, a number of specific tests were conducted, including several that, to our knowledge, have either never been previously conducted (such as multi-KW repetition-rate operation) or have not been performed since the early 1990s (direct IPPT thrust measurements).2 Conical theta-pinch IPPT thrust stand measurements are presented in Fig. 1 while various time-integrated and time
The operation of a 5-msec, quasi-continuous wave (CW) 10-Hz repetition-rate pulsed single-frequency Nd:YAG oscillator with a feedback stabilized frequency bandwidth of less than 200 kHz is described. The major problems associated with frequency stabilization for quasi-CW operation, frequency chirp and maintenance of the frequency control signal, are described, and the design of a wide-bandwidth frequency locking system with a search and chirp compensation loop is given. This frequency stabilized oscillator is a potentially useful laser source for atmospheric wind velocity measurements by coherent lidar Doppler velocimetry, which require a laser source with less than 1 MHz linewidth, and also has applications in high-resolution nonlinear spectroscopic studies at Fourier-transform-limited linewidths.
Laser altimetry provides a high-resolution, high-accuracy method for measurement of the elevation and horizontal variability of Earth-surface topography. The basis of the measurement is the timing of the round-trip propagation of short-duration pulses of laser radiation between a spacecraft and the Earth's surface. Vertical resolution of the altimetry measurement is determined primarily by laser pulsewidth, surface-induced spreading in time of the reflected pulse, and the timing precision of the altimeter electronics. With conventional gain-switched pulses from solid-state lasers and sub-nsec resolution electronics, sub-meter vertical range resolution is possible from orbital attitudes of several hundred kilometers. Horizontal resolution is a function of laser beam footprint size at the surface and the spacing between successive laser pulses. Laser divergence angle and altimeter platform height above the surface determine the laser footprint size at the surface, while laser pulse repetition-rate, laser transmitter beam configuration, and altimeter platform velocity determine the space between successive laser pulses. Multiple laser transitters in a singlaltimeter instrument provide across-track and along-track coverage that can be used to construct a range image of the Earth's surface. Other aspects of the multi-beam laser altimeter are discussed.
Inductive pulsed plasma thrusters are spacecraft propulsion devices in which energy is capacitively stored and then discharged through an inductive coil. While these devices have shown promise for operation at high efficiency on a range of propellants, many technical issues remain before they can be used in flight applications. A conical theta-pinch thruster geometry was fabricated and tested to investigate potential improvements in propellant utilization relative to more common, flat-plate planar coil designs. A capacitor charging system is used to permit repetitive discharging of thrusters at multiple cycles per second, with successful testing accomplished at a repetition-rate of 5 Hz at power levels of 0.9, 1.6, and 2.5 kW. The conical theta-pinch thruster geometry was tested at cone angles of 20deg, 38deg, and 60deg, with single-pulse operation at 500 J/pulse and repetitionrate operation with the 38deg model quantified through direct thrust measurement using a hanging pendulum thrust stand. A long-lifetime valve was designed and fabricated, and initial testing was performed to measure the valve response and quantify the leak rate at beginning-of-life. Subscale design and testing of a capacitor charging system required for operation on a spacecraft is reported, providing insights into the types of components needed in the circuit topology employed. On a spacecraft, this system would accept as input a lower voltage from the spacecraft DC bus and boost the output to the high voltage required to charge the capacitors of the thruster.