Showing papers in "Journal of Spacecraft and Rockets in 1976"
TL;DR: In this paper, a three-axis closed-loop attitude control system for earth-observatory momentum bias spacecraft is proposed, where a horizon scanner and magnetometer are employed to measure attitude errors; no yaw sensor is required.
Abstract: A three-axis closed-loop attitude-control system for earth-observatory momentum-bias spacecraft is proposed. A horizon scanner and magnetometer are employed to measure attitude errors; no yaw sensor is required. Appropriate control signals are generated and used to command variable-strength electromagnets, which interact with the geomagnetic field to reduce observed errors. This system provides for initial acquisition, precession control, nutation damping, and pitch-axis momentum control. All functions are performed autonomously - no ground-station interaction is required. Analytical expressions predicting system response are compared with numerical solutions of the governing equations, and with the results of the application of Floquet Theory.-
258 citations
TL;DR: In this paper, a computer algorithm for extracting a quaternion from a direction-cosine matrix (DCM) is described, which provides a four-parameter representation of rotation.
Abstract: A computer algorithm for extracting a quaternion from a direction-cosine matrix (DCM) is described. The quaternion provides a four-parameter representation of rotation, as against the nine-parameter representation afforded by a DCM. Commanded attitude in space shuttle steering is conveniently computed by DCM, while actual attitude is computed most compactly as a quaternion, as is attitude error. The unit length of the rotation quaternion, and interchangeable of a quaternion and its negative, are used to advantage in the extraction algorithm. Protection of the algorithm against square root failure and division overflow are considered. Necessary and sufficient conditions for handling the rotation vector element of largest magnitude are discussed
105 citations
TL;DR: In this article, a performance level compatible with 30% thrust efficiency at 1500 s specific impulse at 1 mlb (4.45 mN) of thrust was demonstrated with the parallel rail pulsed ablative thruster using Teflon propellant.
Abstract: A performance level compatible with 30% thrust efficiency at 1500 s specific impulse at 1 mlb (4.45 mN) of thrust was demonstrated with the parallel rail pulsed ablative thruster using Teflon propellant. Parametric variations of interelectrod e spacing and included angle were performed. In addition, Teflon was replaced by other thermoplastics and was also seeded with 10% and 30% LiOH and InBr in an evaluation of alternative propellants. Both the conventional breech-fed and later side-feed electrode/propellant configurations were tested. With the same initial conditions it was shown that the breech-fed geometry yields higher thrust efficiency than the side-fed geometry because of the higher specific impulse generated (i.e., up to 5300 s using Teflon). Results of parametric studies indicate that, for high thrust/power and moderately high specific impulse, virgin Teflon propellant with an interelectrode spacing of 3.0 in. (7.62 cm) and zero degree interelectrode included angle in the side-fed configuration are best. A simple semiempirical analytic model is presented where it is shown that the broad range in performance characteristic of this device is related to the degree to which magnetoplasmadynamic and ordinary gasdynamic acceleration mechanisms can be made more or less dominant in the acceleration of the ablated propellant.
74 citations
TL;DR: In this article, a candidate encounter sequence is presented for an equatorial tour of the Galilean moons and a new jargon is introduced in discussions to describe the exploitation of gravity-assist trajectories using the giant satellites for orbit control.
Abstract: Advanced mission and orbit planning efforts are currently in progress for a Mariner-class Jupiter orbiter. Baseline spacecraft and orbit design criteria are the goals of a NASA effort to define such a mission. Orbit design concepts that have been discovered during the early stages of mission planning are both challenging and exciting. A description is given of several such concepts that may greatly increase the flexibility and scientific return of orbiters designed for close study of the Galilean satellites and exploration of the Jovian system. Some new jargon is introduced in discussions to describe the exploitation of gravity-assist trajectories using the giant satellites for orbit control. Orbit 'pumping' and 'cranking' and 'resonance hopping' are defined and shown to be dynamically feasible means of controlling the orbit and, thus, the scientific return. A candidate encounter sequence is presented for an equatorial tour of the Galilean moons.
56 citations
TL;DR: In this article, an experimental investigation was conducted to determine the effect of a uniform pitching motion on a slender axisymmetric body while undergoing large excursions in angle of attack.
Abstract: : An experimental investigation was conducted to determine the effect of a uniform pitching motion on a slender axisymmetric body while undergoing large excursions in angle of attack Force and moment measurements were obtained for a slender tangent-ogive/cylindrical body over a range of Reynolds numbers from 50,000 to 140,000 while varying the angle of attack from zero to 90 degrees and the pitch rate between zero and 281 degrees per second Smoke flow visualization studies were used as an aid in assessing wake vortex transitions
49 citations
TL;DR: In this article, a linear regression model is used to establish a quantitative description of side forces on bodies of revolution at high incidence and zero side slip, and a data base is assembled concerning the key side force characteristics of maximum measured side force, angle of attack at which it occurs, and minimum angle of incidence at which a side force is observed (onset angle).
Abstract: Linear regression techniques are used to establish a quantitative description of side forces on bodies of revolution at high incidence and zero side slip. A data base is assembled concerning the key side force characteristics of maximum measured side force, angle of attack at which it occurs, and minimum angle of incidence at which a side force is observed (onset angle). This information is examined to determine the important trends, and a linear regression model is developed for these quantities to include only those variables which are statistically significant. Results indicate that the peak side force coefficient is a function of Mach number, and only slightly of Reynolds number. Nose fineness is the critical model dimension which suggests that peak side forces are a product of the nose flowfield. Blunting of the nose reduces the magnitude of the side force but the degree of bluntness does not appear to be important. The angle at which the maximum side force occurs is found to be dependent on model length and Mach number, while the onset angle is a function of model length only.
45 citations
TL;DR: The paper examines the validity of the assumption that certain appendage-distributed (modal) coordinates can be truncated from a system model without unacceptable degradation of fidelity in hybrid coordinate dynamic analysis for attitude control of spacecraft with flexible appendages.
Abstract: The paper examines the validity of the assumption that certain appendage-distributed (modal) coordinates can be truncated from a system model without unacceptable degradation of fidelity in hybrid coordinate dynamic analysis for attitude control of spacecraft with flexible appendages. Alternative truncation criteria are proposed and their interrelationships defined. Particular attention is given to truncation criteria based on eigenvalues, eigenvectors, and controllability and observability. No definitive resolution of the problem is advanced, and exhaustive study is required to obtain ultimate truncation criteria.
38 citations
34 citations
31 citations
31 citations
TL;DR: In this paper, a fine-mesh screen, compartmented tank system was developed for the Orbiter RCS application, which provides the required expulsion under widely differing highg boost abort and reentry vectors oriented 119 deg apart and during on-orbit operation under omnidirectional low-g conditions.
Abstract: A program was conducted to provide the technology base for the SS/RCS flight tankage. Through a combination of analysis, subscale testing and computer predictions, a surface tension acquisition/expulsion system design was developed for the Orbiter RCS application. A full-scale tank system was fabricated and ground verification testing was conducted. Cleaning, inspection, fill and drain, and one-g expulsion performance were demonstrated. Results show that the fine-mesh screen, compartmented tank system provides the performance, flexibility, reusability, and other characteristics required by the pulsing, high flowrate RCS. It provides the required expulsion under widely differing high-g boost abort and reentry vectors oriented 119 deg apart and during on-orbit operation under omnidirectional low-g conditions.
TL;DR: In this article, the effects of geomagnetic substorms on the design of spacecraft, especially questions relating to the choice of surface conductivities, concepts of spacecraft •'ground," and shielding requirements are also discussed.
Abstract: Large potential differences and accompanying arc discharges between various parts of a spacecraft may arise as a result of geomagnetic substorms. A large number of electronic malfunctions and otner miscellaneous aberrations in the operation of spacecraft subsystems have been attributed to the substorm phenomenon. This paper reviews some of the substorm related operational problems experienced by synchronous altitude spacecraft systems, discusses the interaction of spacecraft with the substorm plasma environment, and presents the results of spacecraft charging model calculations. The effects of geomagnetic substorms on the design of spacecraft, especially questions relating to the choice of surface conductivities, concepts of spacecraft •'ground," and shielding requirements are also discussed.
TL;DR: In this article, a software technique for onboard detection and identification of hard failures and leaks of the shuttle orbiter reaction control subsystem jets during the orbital flight phase is presented, using only the gimbal angle and linear velocity measurements available from the orbiter inertia! measurement unit.
Abstract: A software technique for onboard detection and identification of hard failures and leaks of the shuttle orbiter reaction control subsystem jets during the orbital flight phase is presented. The method uses only the gimbal angle and linear velocity measurements available from the orbiter inertia! measurement unit. The system produces rotational state estimates required by the attitude autopilot in addition to performing failure identification. Uncoupled steady-state constant covariance extended Kalman filters with residual traps are employed for rotational and translational state estimation, and generalized likelihood ratio tests are made for failure identification. Rigid body simulation results indicate station-level identification times of less than 2 sec for primary jet hard failures and less than 70 sec for vernier jet hard failures and primary jet leaks.
TL;DR: Aerodynamic properties of a Yawed Cylinder at a Mach Number of 4.15 and high Reynolds number are described in this paper. But the results of the analysis are limited.
Abstract: Penland, J.A., "Aerodynamic Characteristics of a Circular Cylinder at Mach Number 6.86 and Angles-of-Attack up to 90°," NACA, RM L54A14, Washington, D.C., March 1954. Dewey, C.F., "Near Wake of a Blunt Body at Hypersonic Speeds," AIAA Journal, Vol. 3, June 1956, pp. 1001-1010. Beckwith, I.E. and Gallagher, J.J., "Local Heat Transfer and Recovery Temperature on a Yawed Cylinder at a Mach Number of 4.15 and High Reynolds Number," NASA TR R-104, Washington , D.C.,1961. Bertin, J.J., Lamb, J.P., Zickler, J.L., and Goodrich, W.D., "Flow -Field Measurements for Space-Shuttle Related Cylindrical Configurations in Hypersonic Streams," AIAA Paper 72-294, San Antonio, Texas, 1972; see also "Heat Transfer Meaurements for Cylindrical Configurations in Hypersonic Streams," Journal of Spacecraft and Rockets, Vol. 10, March 1973, pp. 217-218. 7 Chapman, D.R., "A Theoretical Analysis of Heat Transfer in Regions of Separated Flow," NACA, TN 3792, Washington, D.C., Oct. 1956.
TL;DR: Two Kalman filtering integration techniques are presented for using gradiometer measurements to compensate an inertial navigation system for gravity disturbances and one appears to offer significant theoretical and practical advantages.
Abstract: Two Kalman filtering integration techniques are presented for using gradiometer measurements to compensate an inertial navigation system for gravity disturbances. Both approaches use knowledge of the reference ellipsoid gravity field in addition to gradiometer data and lead to bounded residual gravity-induced navigation errors. Although the two compensation techniques are similar, one appears to offer significant theoretical and practical advantages. The mechanization and error equations for each technique are outlined and the significance of various terms which appear in these equations is discussed in the context of gradiometer-aided inertial navigation.
TL;DR: In this article, the radiative, heating-rate distributions about spherically capped, 60 deg half-angle conical entry probes to the outer planets Jupiter, Saturn, and Uranus were defined.
Abstract: A study has been conducted to define the radiative, heating-rate distributions about spherically capped, 60 deg half-angle conical entry probes to the outer planets Jupiter, Saturn, and Uranus. The radiative heating rates are calculated by a direct method for the solution of the inviscid flow equations with coupled, nongray radiative transport. Results are presented for a range of entry conditions using the atmospheric models denoted as nominal, warm, and cool for each planet which show that both the magnitude and distribution of radiative heating rate are strongly dependent on the atmospheric model and the free-stream conditions.
TL;DR: In this paper, Schapery has generalized the Barenblatt model to the viscoelastic case and thus developed a model that appears to be consistent with both theory and experiment.
Abstract: The analysis of time-dependent crack propagation in viscoelastic materials in general, and solid propellants in particular, has been hampered by the difficulty of the mathematical analysis of a cracked viscoelastic material. The viscoelastic solution to the singular line crack has not provided a realistic fracture criterion, whereas various approximations to this problem have compared more favorably with experiments. However, it is difficult to assess the generality of the approximations involved. Recently, Schapery has generalized the Barenblatt model to the viscoelastic case and thus developed a model that appears to be consistent with both theory and experiment. The objective of the present work was to obtain a detailed comparison of this model with laboratory results available in the literature on a PBAN solid propellant. The results of the comparison are extremely good over a wide range of variables. A time-dependent fracture energy is found to result which can be incorporated readily into the theory.
TL;DR: In this article, a combined gravity gradiometer-inertial navigation system integrated for correction of the gravity anomaly will also measure and hence can be used to map the gravitational anomaly field on a real time basis.
Abstract: Theme R there has been considerable interest in the possibility of correcting high performance inertial navigation systems for gravitational anomaly induced errors using gravity gradient measurements aboard the moving vehicle. A combined gravity gradiometer-inertial navigation system integrated for correction of the gravity anomaly will also measure and hence can be used to map the gravitational anomaly field on a real time basis. The primary purpose of this paper is to evaluate the potential of obtaining relative position updates with such a system by cross-correlating the measured to previously mapped gravity gradients over a known location. Both gravity gradient and gravity anomaly map matching have been analyzed to compare their relative merit. To restrict the scope of the problem, the analysis has been conducted with one element of the gravity gradient tensor in a one-dimensional cross correlation process. The accuracy of relative position updates are presented for various vehicle speeds and for a range of geodetic instrumentation performance levels.
TL;DR: In this paper, a significant increase in damping effectiveness over simple dynamic coupling is achieved through integration of a simple rotor-mounted nutation sensing accelerometer with the despun payload control system.
Abstract: Substantial active nutation control of a dual-spin spacecraft is achieved through augmentation of the fundamental dynamic coupling between the payload azimuth pointing control system and the vehicle nutation dynamics. A significant increase in damping effectiveness over simple dynamic coupling is achieved through integration of a simple rotor-mounted nutation sensing accelerometer with the despun payload control system. This technique is shown to overcome limitations inherent in both passive mechanical damping systems and active nutation control through simple dynamic coupling. The system dynamics and the control laws of the composite nutation and pointing control systems are developed. A closed-form expression for the resulting damping strength is derived. Implementation and functional operation of the control system are discussed. Analysis of the characteristics of the interaction provides design tools for optimization of the system dynamics to achieve maximum effectiveness. A design example is presented for illustration, with simulation results that verify the analytical conclusions.
TL;DR: In this article, the camera system described in this paper was subjected to an extensive test and calibration program prior to launch and subsequent analyses, absolute photometric accuracies of 8.0% may be achieved.
Abstract: Two Viking spacecraft each consisting of an Orbiter and a Lander are on trajectories toward Mars with arrival dates in June and August 1976. A Visual Imaging Subsystem consisting of two slow-scan television cameras forms part of the scientific payload of each Orbiter. These cameras will be used to evaluate the potential landing sites on Mars and to conduct other scientific investigations of the planet. The camera system described in this paper was subjected to an extensive test and calibration program prior to launch. Based on this calibration and subsequent analyses, absolute photometric accuracies of 8.0% may be achieved. Surface resolution exceeding 100 meters will be achieved from the periapsis portion of the Viking orbits. The inherent geometric accuracies of the Orbiter cameras supersede those of previous planetary missions. The analyses of images acquired during the cruise phase of the mission confirms that the cameras have survived the rigors of launch and are performing in a manner consistent with the prelaunch calibrations.
TL;DR: The economics and development program for the OSPP are briefly discussed in this article, where photovoltaic conversion, baseline configuration, microwave reception and power generation, transportation, assembly, and maintenance are considered.
Abstract: Orbital solar power plants are considered from the viewpoint of technology and costs in terms of photovoltaic conversion, baseline configuration, microwave reception and power generation, transportation, assembly, and maintenance. The economics and development program for the OSPP are briefly discussed. (WDM)
TL;DR: In this article, the minimum propellant optimal rendezvous maneuver of two cosmic vehicles on circular and elliptical orbits is examined, where the target without propulsion moves on an orbit around the Earth or other planets as a satellite, and the tracking vehicle with variable thrust propulsion moving on a close orbit.
Abstract: The minimum propellant optimal rendezvous maneuver of two cosmic vehicles on circular and elliptical orbits is examined. It is assumed that the target without propulsion moves on an orbit around the Earth or other planets as a satellite, and the tracking vehicle with variable thrust propulsion moves on a close orbit. The problem of the determination of the optimal laws of variation which characterize the minimum propellant orbital rendezvous is formulated as an extremum variational problem with constraints, where the linear motion equations of the tracking vehicle are considered. On the basis of the optimal laws established in this manner numerical applications of practical interest are carried out.
TL;DR: In this article, a preliminary design study of a 1.2m, cooled infrared telescope for use on 7 to 28 day Spacelab missions is presented. But it is not clear whether the infrared telescope will be able to see more than 30 times deeper into the universe than is now possible.
Abstract: NASA has completed a preliminary design study of a 1.2-m, cooled infrared telescope for use on 7 to 28 day Spacelab missions. Beryllium optics, cooled to below 20° K with supercritical helium, are used. Noise in the infrared telescope will be less, by a factor of 10 , than that of ground-based telescopes operating at 10 /*m (when the telescopes are operated over comparable bandwidths and fields of view). The Spacelab infrared telescope facility will enable astronomers to see more than 30 times deeper into the universe than is now possible over most of its 5-200 jim optimum spectral range. Such capability will initiate a spectacular advance in our knowledge and understanding of the cool regions of the universe where molecules and dust are the predominant radiators and absorbers.
TL;DR: In this paper, a mercury ion thruster was developed for efficient operation at the nominal 1-mlb thrust level with a specific impulse of about 3,000 sec and a total power consumption of about 120 W. At a beam voltage of 1,200 V and beam current of 72 mA, the discharge chamber operates with a propellant efficiency of 93.8% at an ion-generation energy of 276 eV/ion.
Abstract: A mercury ion thruster has been developed for efficient operation at the nominal 1-mlb thrust level with a specific impulse of about 3,000 sec and a total power consumption of about 120 W. At a beam voltage of 1,200 V and beam current of 72 mA, the discharge chamber operates with a propellant efficiency of 93.8% at an ion-generation energy of 276 eV/ion. The 8-cm diameter thruster advances proven component technology to assure the capability for thruster operation over an accumulated beam-on time in excess of 20,000 hours with a capability for 10,000 on-off duty cycles. Discharge chamber optimization has combined stable current-voltage characteristics with high performance efficiency by careful placement of the discharge cathode near the location of a magnetic-field zero just upstream of the thruster endplate.
TL;DR: In this paper, the results of design and development activities and the fire testing of fast response (0.005 sec), high performance (300 sec 75p), 5 IbF bipropellant (N2O4/MMH) engines to obtain performance and life durability data are presented.
Abstract: The subject program advanced the technology for high-performance, rapid response, long-lived bipropellant engines in the 5-lb. thrust class. The results of design and development activities and the fire testing of fast response (0.005 sec), high-performance (300 sec 75p), 5 IbF bipropellant (N2O4/MMH) engines to obtain performance and life durability data are presented. Data were obtained from over 17,000 sec of firing duration and over 400,000 engine starts. Steady-state and pulsing performance, and engine response data were obtained over a range of tank pressures and at environmental temperatures ranging from 20-120° F. Thermal characteristics and heat rejection rates were experimentally evaluated as injector and chamber designs were varied. Simulated duty cycles were demonstrated with three engines. These data were utilized in the forecast of engine life and reliability and the assessment of exhaust plume contamination. Background P RIOR to 1963, few spacecraft missions required the injection of a payload into orbit. Furthermore, there was little need or room for an onboard propulsion system as part of an orbiting package. During the period 1963 to 1968, payload weights increased and a need for station-keeping developed as mission goals became more ambitious. Reaction control systems employed catalytic decomposition of hydrogen peroxide and/or cold gas jets. The instability of the hydrogen peroxide under storage and the need for pressure relief valves made the reliability of this system inherently low.
TL;DR: A technique is described for theoretical, statistical evaluation of the thrust imbalance of pairs of solid-propellant rocket motors firing in parallel, and sets of the significant variables are selected using a random sampling technique and the imbalance calculated for a large number of motor pairs.
Abstract: A technique is described for theoretical, statistical evaluation of the thrust imbalance of pairs of solid-propellant rocket motors (SRMs) firing in parallel. Sets of the significant variables, determined as a part of the research, are selected using a random sampling technique and the imbalance calculated for a large number of motor pairs. The performance model is upgraded to include the effects of statistical variations in the ovality and alignment of the motor case and mandrel. Effects of cross-correlations of variables are minimized by selecting for the most part completely independent input variables, over forty in number. The imbalance is evaluated in terms of six time - varying parameters as well as eleven single valued ones which themselves are subject to statistical analysis. A sample study of the thrust imbalance of 50 pairs of 146 in. dia. SRMs of the type to be used on the space shuttle is presented. The FORTRAN IV computer program of the analysis and complete instructions for its use are included. Performance computation time for one pair of SRMs is approximately 35 seconds on the IBM 370/155 using the FORTRAN H compiler.
TL;DR: In this paper, the authors compared the impact of surface roughness on the performance of a delta wing shuttle orbiter with and without longitudinal pressure gradients on two-dimensional zero-pressure gradient flows and three-dimensional flows.
Abstract: Tripping effectiveness of surface roughness on a delta wing shuttle orbiter model at 20 deg angle of attack is compared to that on plane and axisymmetrical bodies with and without longitudinal pressure gradients. The experimental data presented are compared on the basis of effective roughness Reynolds number since this parameter is not sensitive to flow conditions downstream of the roughness. The discussion covers the effective roughness Reynolds number as a function of roughness position Reynolds number, effective size ratio as a function of pressure gradient and distance from vehicle nose, and effect of spanwise roughness position on roughness effectiveness. It is shown that conventional criteria for sizing roughness elements which promote transition in two-dimensional zero-pressure gradient flows are insufficient for high-pressure gradient flows and three-dimensional flows. Roughness much smaller than that given by conventional criteria can cause transition and significantly increase the heating load.
TL;DR: The relationship between the longitudinal mode oscillatory combustion phenomena in solid-propellant motors and the motor configuration was experimentally examined using small test motors with various grain configurations.
Abstract: 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.
TL;DR: In this paper, a quasi-steady analytical theory is developed to investigate the effects that unsymmetrical stability derivative characteristics can have on the transient angular motion behavior of rolling re-rentry vehicles.
Abstract: A quasi-steady analytical theory is developed to investigate the effects that unsymmetrical stability derivative characteristics can have on the transient angular motion behavior of slender rolling re-rentry vehicles. The theoretical results indicate that the presence of unsymmetrical stability derivatives can alter the motion patterns, frequency, and damping sufficiently to invalidate the use of previous analytical models developed for vehicles with symmetrical stability derivative characteristics. Between the resonances small asymmetries in the static moment slope coefficients are shown to cause a static instability which results in an exponential growth of the transient angle of attack. The analytically derived expressions and the analytically predicted trends are verified by numerical analyses.