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A hybrid rockoon, which is a hybrid rocket launched from a high-altitude balloon, is proposed. The rocket system configuration was studied for single-stage or multiple-stage rockets, and the perfor...

In this paper, design and optimization approaches are applied to a hybrid-powered small-satellite launcher, namely a deterministic and a robust-based method. A unique hybrid rocket engine is consid...

The energetic electrostatic analyzer has been developed to act as a miniaturized, rugged, and low-cost instrument capable of providing in situ measurements of ion or electron energy and subsequent ...

A feasibility study related to the scale-up of solid fuel with multiport was performed using the regression rate equation derived from a laboratory-scale hydrogen peroxide (H2O2) hybrid rocket usin...

The Boundary Layer Transition (BOLT) Flight Experiment is to launch soon on a sounding rocket to obtain hypersonic data on a complex geometry with swept leading edges and concave surfaces. BOLT was...

This paper describes three-dimensional (3D) stereo reconstruction efforts for the full-scale supersonic parachute deployment tests performed under the NASA Advanced Supersonic Parachute Inflation R...

The Advanced Supersonic Parachute Inflation Research Experiment (ASPIRE) was a series of sounding rocket flights aimed at understanding the dynamics of supersonic parachutes that are used for Mars ...

Accepted: 2018-05-21, 資料番号: SA1180213000

The Supersonic Flight Dynamics Test was a full-scale flight test of aerodynamic decelerator technologies developed by the Low-Density Supersonic Decelerator technology development project. The purp...

On 8 November 2016 at 1030 hrs, the Hybrid Experimental Rocket Stuttgart (HEROS) 3 was launched from the European Space and Sounding Rocket Range (ESRANGE) Space Center to an apogee altitude of 32,...

Accepted: 2015-08-20, 資料番号: SA1150226000

Accepted: 2014-03-27, 資料番号: SA1150043000

The primary objective of the Radio Aurora Explorer mission is to study plasma instabilities that lead to magnetic field aligned irregularities of electron density in the lower polar ionosphere (80–...

The mission architecture of SHEFEX II features a two-stage solid propellant sounding rocket vehicle on a suppressed trajectory, which is induced by a cold gas pointing maneuver of the vehicle before second stage ignition. The impact point is subject to a 3-� dispersion of roughly�110 kmin downrange and�90 kmin crossrange, which makes a recovery of the vehicle particularly difficult, as the whole impact area is located off shore and the vehicle needs to be recovered by ship. As the major part to dispersion is contributed during the atmospheric ascent of the vehicle, a control algorithm is developed that considers the actual deviation from the nominal trajectory after atmospheric exit and recommends a vehicle pointing that corrects for this deviation. The analytic control algorithm is found by linear/quadratic approximation of the impact point sensitivity towards the deviations after atmospheric exit and to the pointing angles. The effectiveness of the algorithm is tested by implementing it in a full six-degree-offreedom simulation and applying dispersion factors in a Monte Carlo simulation. The result is a reduction of the impact point dispersion area by about 78%.

A low cost method for reducing the dispersion in the trajectory of an unguided, spinstabilized, sounding rocket is developed and presented. The method is particularly suited to scramjet flight experimentation because the approach increases the likelihood of meeting Mach number and dynamic pressure objectives. The paper discusses the design and model of the scramjet payload, and two-stage launch vehicle, and the nominal trajectory, as well as a Monte Carlo analysis to quantify the likelihood of a successful scramjet test. Using the results of this analysis, a method is presented for reducing the dispersion in freestream conditions during the scramjet test window. The dispersion reduction is accomplished by modifying the time delay between the burnout of the first stage booster and the ignition of the second stage based on the vehicle state measured during the interstage coast. This method increases the likelihood of a successful test from 71% to 99% without adversely affecting range safety. Since the design and implementation of a vehicle guided control system is not required, this method is relatively inexpensive, making its use highly desirable for low cost scramjet flight experimentation.

This paper derives approximate angle-of-attack statistics suitable for boost-phase structural loading estimates on unguided, fin-stabilized sounding rockets. Sounding rockets are assumed to lift off with a large constant vertical acceleration.Theirrigid-bodyrotationsaremodeledasundampedshort-period motionswithoutanyvelocityvector rotation;the onlysource ofpitch/yawtorquesisviaaerodynamic staticstability. Theverticalacceleration causesthe dynamic pressure to increase rapidly, leading to time-varying coefficients in the short-period equations and, therefore, gust responses that require nonstationary analyses. Transforming the independent variable from time to altitude enables calculation of a simple lateral velocity sinusoidal gust impulse response function. Next, the total response for a single instantiation is found by superposition of all its gust impulses. Then, convolution to find the variance in transverse velocity is found based on the Dryden gust autocorrelations. A closed-form result for the standard deviation in boost-phase gust angle of attack is obtained and compared with both its high-altitude asymptote and the classical sharp-edged (step function) gust response. At altitudes above about two pitch wavelengths, the asymptote provides an accurate result, whereas the classical sharp-edged gust model significantly underestimates the gust response, except for regions very near the ground.

A compact and rugged fiber-coupled liquid volume sensor designed fo r flight on a sounding rocket platform is presented. The sensor consists of a Mach-Zehnder interferometer capable of measuring the amount of liquid contained in a tank under any gravitational conditions, including a microgravity environment, by detecting small changes in the index of refraction of the gas contained within a sensing region. By monitoring changes in the interference fringe pattern as the system undergoes a small compression provided by a piston, the ullage volume of a tank can be directly measured allowing for a determination of the liquid volume. To demonstrate the technique, data are acquired using two tanks containing different volumes of liquid, which are representative of the levels of liquid in a tank at different time p eriods during a mission. The two tanks are independently exposed to the measurement apparatus, allowing for a determination of the liquid level in each. In a controlled, laboratory test of the unit, the system demonstrated a capability of measuring a liquid level in an individual tank of 10.53 mL with a 2% error. The overall random uncertainty for the flight system is higher than that one test, at ±1.5 mL.

This paper presents trajectory reconstruction of the ST-9 (space technology) sounding rocket experiment using the onboard inertial measurement unit data and descent imagery. The raw inertial measurement unit accelerometer measurements are first converted into inertial acceleration and then used in trajectory integration. The descent images are preprocessed using a map-matching algorithm and unique landmarks for each image are created. Using the converted inertialmeasurement unit data anddescent images, the result fromdead-reckoning and the kinematicfix approaches are first compared with the global positioning system measurements. Then, both the inertial measurement unit data and landmarks are processed together using a batch least-squares filter and the position, velocity, stochastic acceleration, and camera orientation of each image are estimated. The reconstructed trajectory is compared with the global positioning system data and the corresponding formal uncertainties are presented. The result shows that inertial measurement unit data and descent images processed with a batch filter algorithm provide the trajectory accuracy required for pinpoint landing.

the device as measured from sounding rocket experiments conducted by the U.S. Air Force Research Laboratory, Space Vehicles Directorate. A new modeling approach to predict the internal acoustic response particular to sounding rockets is presented wherein the interior noise results from the time varying accelerations acting through the forward bulkhead. This model is different from typical approaches used for payload fairing noise prediction in which the primary noise sources are the rocket motors. This model would also apply to rockets of a similar aspect ratio and acceleration profile. The active acoustic absorber presented in this work can be tuned for optimal performance just minutes before launch. Acceleration and acoustic measurements from two sounding rocket launchesarepresentedtovalidate themodelingapproachandtodemonstrate theperformanceoftheactiveacoustic absorber.Datashowedthatasingledeviceachievedan8.8dBreductioninthesoundpressurelevelfrom20to300Hz.

The direct simulation Monte Carlo method is used to model a transient stage separation of a generic sounding rocket at 100 km. Lower-stage movement is included, and the flow and surface properties are simulated during the first second after thruster ignition. Both liquid and solid propellant thrusters are examined with a thrust of 25 and 34 kN, respectively. Four different simulation scenarios are considered that allow analysis of the impact of the stage motion, explicitly including unsteady flow effects. Unsteady flow effects are small enough that quasi-steady-state modeling appears to be adequate for this general staging scenario. The influence of direct simulation Monte Carlo statistical fluctuations on the stage trajectory is insignificant comparedwith the total contribution of the plume force. We also examine the radiation environment, including the plume–atmosphere shock and plume–lower-stage impingement.

Over the past several decades, atomic oxygen measurements taken from sounding rocket sensor payloads in the altitude range of 80-140 kilometers have shown marked variability. Many sounding rocket payloads contain atomic oxygen sensors that are located in close proximity to the payload surface, and are thus significantly influenced by flow field disturbances. Although several additional factors including Doppler shift and sensor contamination may also play a significant role in the accurate measurement of atomic oxygen concentrations, this work focuses solely on the effects due to the flow field. The present study utilizes the three-dimensional, steady-state, direct simulation Monte Carlo technique. In addition, the lower altitudes corresponding to near-continuum flow are solved via the Navier-Stokes equations with slip wall boundary conditions. The flow is simulated at 13 different altitudes, each with three separate rocket orientations, along both the rocket's upleg and downleg trajectory for a total of 75 simulations. The numerical simulations show conclusively that the relative magnitudes of undisturbed versus disturbed atomic oxygen concentrations are highly dependent upon rocket orientation, and provide a quantitative means by which existing atomic oxygen concentration data sets may be corrected for aerodynamic influences.

The tip-tilt mirror (TTM) system was developed for the extreme ultraviolet (XUV) Cassegrain telescope aboard the Institute of Space and Astronautical Science sounding rocket. The spatial resolution of the telescope is about 5 arcsec, whereas the rocket pointing is only controlled to be within ±0.5 deg of the target (sun) without additional stability control. To stabilize the XUV image within about 5 arcsec on the focal plane, the TTM system controls the tilt of the secondary mirror with two-axis fixed-coil magnetic actuators. The TTM system has a wide tilt angle and can drive the large secondary mirror at high frequency. The two position-sensitive detectors, one placed in the telescope and the other in the TTM mechanical structure, are used for closed-loop control of the TTM. The closed-loop control system, which has command and telemetry, is executed by the flight software on the digital signal processor. The TTM has a launch-lock mechanism to protect against launch vibrations up to about 16G. The sounding rocket was launched from the Kagoshima Space Center on 31 January 1998. The TTM worked perfectly during the flight and achieved better than the expected 5-arcsec stability on the focal plane during CCD charge-coupled device exposures.

We report on data from particle instruments making in situ sounding rocket measurements of the particle environment within and near an aluminum plasma jet caused by an explosion in the auroral ionosphere. The Active Plasma Experiment sounding rocket was launched to an altitude of 350 km from the Poker Flat Research Range in January 1999. The payload separated after launch into observing payloads and two explosive plasma jet generators. During the flight, the two explosive packages were detonated, and the observing payloads studied the surrounding environment. The particle instruments measured the resulting plasma jet from a distance of approximately 500 m. The instruments measured ions from 10 to 420 eV, electrons from 10 to 6000 eV, and low-energy electrons from 2 to 1200 eV. After each explosion, the particle instruments recorded the passage of a burst of material past the spacecraft. Analysis of mass-dependent effects, plasma β, and critical ionization velocity parameters are presented, together with a comparison to earlier experimental observations. In particular we note that the duration of the enhanced ion fluxes is controlled by the jet velocity and drops sharply when the jet velocity falls below the critical ionization velocity for each ion, with the peak ion fluxes only observed while v j e t > V c r i t (O+).

This paper discusses the evolution of the American Rocket Society and its experimental programs from the early liquid rocket launches, through static firings of engines, to the experimental aerodynamic development of rocket shapes. The American Rocket Society was founded in 1930 as the American Interplanetary Society by a small group of science-fiction writers who were intrigued by the concept of interplanetary travel or space flight. A short time after its forming, the interplanetary society had become a more down-to-Earth organization. Within a matter of months, a "program of research on the rocket and its possibilities" had been undertaken. Out of this initiative and a trip by one of the founding members to visit the German Interplanetary Society, the Verein fur Rauinschiffahrt E. V. Geschaftsstelle Berlin in 1931, the experimental committee of the American Interplanetary Society was born. The experimental committee would develop five rockets, of which three would actually fly. Their most important contribution to rocketry at the time was the development of an experimental test stand, later used by one of America's first rocket manufacturing companies, Reaction Motors Inc. Four members of the American Rocket Society's experimental committee would be the founders of this company. The experimental phase of the American Interplanetary Society and later the American Rocket Society lasted only about 10 years but led to the advancement of rocket science. American Interplanetary Society's scope had moved far from its origins as group of space enthusiasts to amateur experimenters to the country's leading technical society for rocketry and jet propulsion.

The inverse method and the numerical optimization technique are developed and implemented for the design of space launch vehicles, which satisfy given heat transfer rate, e neness ratio and internal volume constraints, and show minimum drag characteristics. With the proper specie cation of the target pressure distribution, the inverse method is successfully applied to the design of a body with 21% less drag than the initial shape. Several gradient approximations, shape functions, and approximate analysis methods are utilized for the design of optimum nose fairing shapes with heat e ux and volumeconstraints. Several strategies are implemented to the design examples in hypersonic speeds, including a Korean three-stage sounding rocket. The designed bodies have less drag than the initial bodies while maintaining the surface heat transfer rate at the nose. These methods are demonstrated to be efe cient design tools for the high-speed vehicle design.

A fully reusable rocket vehicle is proposed to demonstrate good operability characteristics both on the ground and in e ight. For achieving technical readiness for future space transportation systems, design considerations not only for higher-performance-related issues but also those for good operability are needed. The proposed vehicle is to be used as a sounding rocket and has the capabilities of ballistic e ight, returning to the launch site, and landing vertically, making use of clustered liquid-hydrogen rocket engines. Before the development of this type of reusable rocket was initiated, a small test vehicle with a liquid-hydrogen rocket engine was built and e ight tested. A demonstration of vertical landing and exerciseof turnaround operation forrepeated e ightsare the major objectivesof the test vehicle. Two e ightswere performed in succession, and the e ight-test operation provided many valuable experiences for designing the fully reusable rocket vehicle.

A high-resolution, low-energy electron spectrometer can be deployed on rockets or satellites to measure absolute spacecraft floating potential relative to true (Earth) ground with an accuracy of ±0.2 V. Features due to the photoionization of nitrogen and atomic oxygen by an extremely sharp solar extreme ultraviolet line (304 A, He II) appear in the electron spectra collected by the instrument. At altitudes where the photoelectron production is local (below 250 km), the peaks that appear in the spectrum are created by electrons of a known kinetic energy: the energy of the ionizing photon minus the energy needed to ionize either atomic oxygen or molecular nitrogen. The spacecraft potential is determined from the apparent shift in the energy of the photoelectron spectral peaks. The instrument is best suited for use in the daytime atmosphere at altitudes between 150 and 250 km. The spectrometer is based on an earlier design, that of the photoelectron spectrometer of the Atmosphere Explorer satellites. Enhancements of the earlier design, including a threefold greater throughput and a specialized scan mode, greatly improve the determination of spacecraft potential. The placement of the instrument on the spacecraft is critical to the success of the measurements.

Microgravity experiments have been performed aboard STS-57, STS-60, and STS-63 with an isothermal hightemperature automated furnace. The equipment for controlled liquid phase sintering experiments-SPACEHAB variant facility was used to conduct powder metallurgy experiments on compacts at temperatures exceeding 1100 ± C. An overview of the evolution of the program from a laboratory mock-up through four sounding rocket e ights is presented, including a description of the hardware that successfully operated during three Space Shuttle missions.Theseaccomplishmentsweretraceddirectlytothedecisionsmadeandnumerouslessonslearnedthroughouttheappliedresearchprogram.Developmentofthehardwarerequiredanextremelyinterdisciplinaryapproach. A continuous path of upgrades was pursued between e ights, increasing capabilities while reducing payload mass and volume.Thisevolutionaryexperiencebaseresultedinsignie cantcostandschedulesavings.Thedesign features of containment and redundancy with the proven e ight performance data minimized expensive component testing and analysis. The hardware was designed with inherent growth capabilities to maximize safety, to promote longterm usage/future e ight opportunities, and to simplify on-orbit operations. The next-generation design scheduled to e y aboard STS-79 (August 1996) also is described.

The Space Physics Experiments Aboard Rockets-3 sounding rocket carried a payload that performed active experiments to diagnose the physical mechanisms and test the effectiveness of several grounding schemes, to study high-voltage bias effects on the performance of solar cells, and to monitor the undisturbed plasma and neutral gas environment of the payload. As a part of that payload, the neutral pressure gauge obtained measurements of the pressure surrounding the payload both during and between the various active experiments. Neutral pressure results show a pressure elevated by as much as two orders of magnitude over atmospheric model-derived pressures for the entire flight. Neutral pressure measurements indicate that gas released during neutral gas (argon) releases, which was one of the mechanisms for grounding the payload, and attitude control system thruster firings (nitrogen) may be displacing the ambient gas, at least from the vantage point of the neutral pressure gauge. The prerelease ambient pressure appears to be the determining factor for whether a gas release causes an increase or decrease in the ambient pressure. Also, the decay time associated with these gas releases, or the time it takes for the measured pressure to return to ambient, is on the order of a few tenths to a few hundredths of a second once the gas valves are shut.

Experimental results from many platforms indicate strong coupling between lightning and the ionosphere and show the necessity of high-frequency sampling in future experiments. We describe the lightning-triggered, burst- mode data acquisition algorithm and circuit used in the NASA sounding rocket 36.111, which is scheduled for launch in 1995. Burst-mode acquisition increases the time resolution by almost two orders of magnitude over that of continuous sampling, but requires a smart controller to select important data. The increase in time resolution will yield sounding-rocket electromagnetic-wave data valid up to the plasma frequency for the first time in the lightning environment. The instrument samples the vector electric field and one component of the magnetic field at 4 MHz with 10-bit, resolution, yielding a burst rate of 320 Mbits/s. Burst-mode instruments use ground commands, precursors, or recorded preburst data to ensure the proper data are recorded. N the early years of space research, the electromagnetic waves launched by lightning strokes were found to disperse in time and to create an audio-frequency radio wave, which whistled when con- verted to sound waves. These whistlers bounce between the Earth's hemispheres along magnetic field lines and are an important loss mechanism for the radiation belts. The study of whistlers has re- vealed much about the plasma around the earth, including the dis- tribution of ionization in the ionosphere.1 Research in the last two decades, including a number of curious ground-based, rocket, satel- lite, and Space Shuttle observations during severe storm conditions, showed that lightning may be an important source of ionospheric disturbance as well. At this time, known interactions between light- ning and the ionospheric plasma are virtually unexplained by exist- ing theories. In situ rocket and satellite measurements performed in the past have been hampered by telemetry bandwidth limitations, and hence, crucial information is absent. A summary of these ex- perimental results is given. The phenomenon of explosive spread F (an event in which the radar backscatter signal due to electron density irregularities from approximately 250 km increases to sizeable levels in tens of mil- liseconds and decays in hundreds of milliseconds) was first reported by Woodman and La Hoz. 2 The NASA rocket flight Thunder Hi (33.022) measured large-amplitude transient electric fields owing to lightning in the ionosphere above an active thunderstorm3 with amplitudes sufficient to drive plasma instabilities in the F region of the ionosphere.4 Woodman and Kudeki5 subsequently conducted a radar experiment that showed explosive spread F is highly correlated with lightning activity. A class of early/fast Trimpi events (a type of transient perturbation in the propagation of subionospheric vlf signals) was reported that cannot be explained by the whistler-based theory of resonant pitch- angle scattering and precipitation of radiation-bel t particles. One theory is that plasma in the mesosphere is heated, resulting in a perturbation to the Earth-ionosphere waveguide, causing the Trimpi event.6'7 The Thunderstorm II sounding rocket (27.120) measured large (10-40 mV/m), long-duration (1 ms) electromagnetic (EM) pulses with electric-field components parallel to the Earth's magnetic field

A sounding-rocket payload to study the generation and emission of electromagnetic waves from a modulated electron beam was developed and successfully flown from the Poker Flat Research Range, Alaska, in March 1992 on a Black Brant 11 sounding rocket. We describe the mission objectives, the instrumentation, flight operations, and preliminary results from the flight. The payload contained a modulated electron gun with a triode arrangement of electrodes allowing modulation of the beam current up to vlf frequencies. Vehicle charging was inhibited by a gas-release system, synchronized with the beam emissions. A network of ground stations was set up to try to receive signals from the modulated beam, and low-light systems were set up to look for evidence of the beam below the payload trajectory. No evidence of beam-induced wave radiation to the ground was detected. However, strong vlf frequencies following the preset program were clearly detected by the diagnostic free flyer and the tethered daughter payload segment. The effect of the gas releases was very marked. The vehicle potential dropped from over 1 kV to about 30 V when the gas was turned on. Some evidence of the light from beam-atmosphere interactions was seen near the end of the flight, but in general, the optical results were not very informative. The launch was successful, mission operations were as planned, and good-quality onboard data were collected throughout the operational part of the flight.

The OEDIPUS-A tethered rocket payload consisted of two spinning subpayloads each with a pair of flexible radial booms and connected by a conductive tether. This first flight of a spinning tethered two-body system achieved a tether deployment of 958 m, which at that time was the longest space tether ever flown. Although the approximately 15-min mission was generally very successful, unexpected dynamic behavior was observed in one of the subpayloads. This paper reviews the attitude dynamics investigation that was undertaken following the flight in an attempt to understand what may have caused this dynamics anomaly. The formulation of a general mathematical model is described that consists of a rigid central body undergoing large rotations with three orthogonal pairs of flexible booms. Tether interaction with the central body is modeled as a forcing term, caused by the tether tension, to the free-body equations. Results of a parametric study are presented, which show that the interaction of the tether with the payload, which was assumed to be negligible prior to the flight, is the likely cause of the unexpected dynamic behavior. Two mechanisms associated with the tether interaction have been identified: one is a forcing term on the payload, which causes a precession of the payload angular momentum vector, and the other is additional energy dissipation in the payload.

Twelve experiments designed to obtain samples and data for a variety of microgravity materials processing investigations were carried on Consort 3, a sounding rocket flight. The overall purpose of the flight was to provide an opportunity to investigate the effects of microgravity on processes with potential for new technology development and commercial applications. The range of experimentati on included investigations of immiscible polymer demixing rates, electrodeposition and electrocodeposition, elastomer-modified epoxy resins, foam formation, a number of biomaterial mixing processes, and polymer thin-film formation. Specific objectives and results are summarized for each experiment. The experiments were contained in a payload module which was launched and recovered within 4 h at the White Sands Missile Range in May, 1990. The microgravity portion of the flight lasted approximately 7 min. Apogee occurred 283 s after liftoff at an altitude of 189 statute miles. The integrated payload module configuration and mission sequence of events are briefly described. Evaluation of flight samples and recorded flight data is discussed for individual experiments. Consort 3 provided a large number of investigators with empirical information on processing materials in microgravity and on the effects of microgravity on the processes performed.

Water outgassing at altitudes greater than 250 km was observed during the Lincoln Laboratory HAVE SLED II series of sounding rocket experiments. The major spectral features include the water vi band at 6.3 /um and rotational emission at wavelengths greater than 15 jim. These data contain information on the rotational and vi vibrational temperatures, as well as the water outgassing rate. An analytical model has been developed to predict the long wavelength infrared spectral radiance from the outgassing of water from a vehicle surface. An outgassing rate of 0.3-3 g/s of H 2O was determined to be consistent with the data. The H2O rotational and vibrational temperatures were estimated to be 220 ± 30 K and 265 ± 30 K, respectively.

During the past decade, the plasma frequency probe (PFP) has evolved into an accurate, proven method of measuring electron density in the ionosphere above about 90 km. The instrument uses an electrically short antenna mounted on a sounding rocket that is immersed in the plasma and notes the frequency where the antenna impedance is large and nonreactive. This frequency is closely related to the plasma frequency, which is a direct function of free electron concentration. The probe uses phase-locked loop technology to follow a changing electron density. Several sections of the plasma frequency probe circuitry are unique, especially the voltage-controlled oscillator that uses both an electronically tuned capacitor and inductor to give the wide tuning range needed for electron density measurements. The results from two recent sounding rocket flights (Thunderstorm II and CRIT II) under vastly different plasma conditions demonstrate the capabilities of the PFP and show the importance of in situ electron density measurements of understanding plasma processes.

The third sounding rocket payload in the Consort program was launched from the White Sands Missile Range on May 16, 1990. It carried 12 experiments designed to investigate materials processes in low gravity. All of the experiments were reflights from the Consort 2 payload that was launched on November 1989 but failed to achieve microgravity because of a malfunction in the launch vehicle. Four national Centers for the Commercial Development of Space participated in the mission. The payload included five experiments and two accelerometer systems that flew on Consort 1 and seven new experiments, designed and developed since Consort 1. Experiments from Consort 1 incorporated hardware modifications and changes in experimental parameters based on mission results. The new experiments covered a variety of polymeric and biological investigations. A new power distribution and control system designed to provide discrete, computer-supervised, experiment power monitoring and control was flight qualified on Consort 3. Consort 3 featured very late access (3-5 h before launch vs 27-30 h for Consort 1) for installation of sensitive biological specimens. The integrated payload and mission sequence of events are described. Changes in the Consort 1 experiments are defined and the objectives, methods, and expectations for new experiments are discussed.

An apparatus that produced the first polyurethane foam in low gravity has been described. The chemicals were mixed together in an apparatus designed for operation in low gravity. Mixing was by means of stirring the chemicals with an electric motor and propeller in a mixing chamber. The apparatus was flown on Consort 1, the first low-gravity materials payload launched by a commercial rocket launch team. The sounding rocket flight produced over 7 min of low gravity during which a polyurethane spheroidal foam of approximately 2300 cu cm was formed. Photographs of the formation of the foam during the flight show the development of the spheroidal form. This begins as a small sphere and grows to approximately a 17-cm-diam spheroid. The apparatus will be flown again on subsequent low-gravity flights.

The CHARGE-2 sounding rocket payload has measured the transient and steady-state charging of a spacecraft in LEO during the emission of a low-power electron beam. The electron beam successfully escaped the emitting spacecraft above 240 km, rather than being degraded by the spacecraft's potentials. These potentials were limited to about half of the 1-kV beam accelerating potential at all latitudes, suggesting that the electron beam was able to escape at altitudes down to 160 km. Electrons created from beam-plasma interactions become increasingly important in the return current below 240 km, and increased with decreasing altitude.

T HE North Star Active Plasma Experiment mission was designed to investigate the dynamics and interactions of highspeed plasma jets in the high-latitude ionosphere. The artificial high-speed plasma jets produced during the North Star mission are relevant to plasma jets produced by electric propulsion thrusters, to electrodynamics associated with spacecraft tethers, and to advance spacecraft charging technologies. In addition, this experiment is relevant to the study of basic plasma physics problems such as momentum coupling between different plasma populations and how this coupling relates to the propagation of Alfven waves and electron acceleration. Specifically, the goals of the North Star mission were to address issues relating to the propagation of high-speed plasma jets across magnetic field lines, investigate the role of the neutral atmosphere in secondary jet ionization processes, and assess the Critical Ionization Velocity hypothesis. The North Star mission1 was launched on 22 January 1999 at 1358:03 UT using a Black Brant XII sounding rocket from Poker Flat, Alaska (Fig. 1). The experiment occurred just after an auroral breakup. The trajectory of the rocket crossed auroral arcs on its ascent prior to the execution of the two active plasma-jet experiments. The plasma jets were produced using a device known as an explosive-type generator (ETG) and resulted in the generation of a high-speed jet of both neutral and ionized aluminum. The neutral aluminum is formed from the rapid recombination of aluminum ions formed at the time of the initial jet initiation. The jets were directed perpendicular to the magnetic field toward instrumented subpayloads located 200–1500 m from the source. The plasma-jet experiments occurred at altitudes of 360 km (ETG-1) and 280 km (ETG-2). A canister of compressed air was released prior to ETG-1 injection to simulate the neutral density at a 150-km altitude in the vicinity of the payload. The North Star plasma-jet papers reviewed in this introduction address diverse topics, such as plasma-jet electrodynamics, plasma-jet magnetic field perturbations, plasma-jet propagation and neutralization, optical signatures of the plasma jet, and the Critical Ionization Velocity (CIV) theory. Erlandson et al.1 present an overview of the North Star mission and a review of the findings, and they discuss the high-speed optical measurements of the high-speed aluminum plasma-jet injection. The emissions were dominated by line emissions caused by neutral aluminum and a continuum thought to be from hot (1500–2000 K), micron-sized debris. The timing from detectors viewing the jet

The current available in situ data on the physical characteristics of the 90-130 km environment depends on sounding rocket and maneuverable satellite systems for its collection. Such data are, of necessity, sporadic in both time and space. The development of tethered systems in Earth orbit will allow global, long-term observations of this environment for the first time. In addition, tethered systems will allow spacial gradiometry at altitudes from 130400 km without the extensive celestial mechanics required to coordinate multiple satellite observing programs. This paper reviews some of the outstanding science questions for the altitude regime from 90-400 km which require the use of tethered observing systems for study. These questions involve the structure, temporal evolution, and composition of the uppermost atmosphere/lowermost ionosphere, the distribution of the high spatial frequency components of the magnetic and gravity fields, and the spatial gradients of the properties of the lowermost magnetosphere/uppermost ionosphere.

The relationship between the longitudinal mode oscillatory combustion phenomena in solid-propellant motors and the motor configuration is experimentally examined using small test motors with various grain configurations. The test results show that the establishment of longitudinal oscillation closely relates to a change of the port area along the motor axis, the internal gas flow conditions, and the motor slenderness. The effects of these parameters on the motor design criterion are qualitatively discussed in relation to the propellant response function and the agglomerative burning of aluminum particles. The test results of an effective mechanical suppression device applied to an actual sounding rocket motor are also presented.

Electrostatic charging of the orbiter has been studied by the Space Shuttle/Spacelab-1 using the Space Experiments with Particle Accelerators (SEPAC). Charging of the orbiter due to electron beam emission has been analyzed using data from a Langmuir probe, floating probes, an electron energy analyzer and a low-light-level TV camera. Charging of the orbiter is found to be strongly dependent upon the attitude of the orbiter with respect to the velocity vector. The orbiter potential has reached the beam acceleration voltage well beyond one kilovolt, which is much higher than that previously observed in electron beam experiments on sounding rockets in the lower ionosphere.

This paper describes a new "integrated method" that allows sounding-rocket development engineers to determine the effects of various perturbing moments on the oscillatory motion of the vehicle. Furthermore, since a pointmass trajectory, aerodynamic pitching and damping coefficients, and an aeroelastic analysis are required for the basic vehicle design, the oscillatory analysis can be performed with very little extra effort. Analysis of a current rocket system by this method simulates the actual flight data very closely.