Showing papers on "Rocket published in 1987"

Collisional interchange instability 2. A comparison of the numerical simulations with the in situ experimental data

Abstract: A two-dimensional model applicable to F region ionosphere plasma instabilities has been developed and described in a companion paper. Here the model is applied to equatorial F region irregularities and is tested against rocket and satellite data. As a diagnostic, simulated data sets are created which are similar to the one-dimensional measurements of plasma density performed by space probes and the Fourier transform of these data is taken in the same manner and used by the space experimentors. Unlike previous simulations of this phenomenon, an inherent anisotropy is found in the instability development which is mirrored in the in situ data. Evidence that the shallow spectral slopes which often characterize spread F rocket spectra near the F peak may be due to a change in the angle between the rocket velocity vector and the characteristic directions in the medium is presented.

Rocket exhaust deflector

Abstract: A missile canister hatch cover is provided which deflects the rocket exhaust from an adjacent uptake channel away from a missile as it leaves the canister. After diverting the uptake exhaust flow away from the missile, the hatch cover is returned to its closed position by the missile rocket exhaust as the missile nozzle clears the canister opening. Overall wear and tear on the launching system from the rocket exhaust is reduced and the uptake flow plume is stopped from expanding above the launcher. During missile launch the hatch cover is unlatched and spring loading opens it more than 90° to a locked position where it interferes with the exhaust flowing from the uptake. The interference deflects the uptake flow away from the missile during flyout to avoid the heating, side thrusts, and contamination associated with the uptake exhaust flow field. After the missile clears the canister and the rocket exhaust begins to impinge on the hatch cover, the hatch cover is unlocked from its open position by actuation of a drag flap that is deployed to help close the hatch cover. An ablative material is applied to exposed surfaces of the hatch cover to prevent heat damage while it is deflecting the uptake exhaust flow.

Results from a series of tethered rocket experiments

Abstract: This paper reports the results of three tethered rocket experiments conducted as part of a U.S.-Japan joint program under way since 1980. The major purpose of the experiments is to obtain technical and scientific data supporting the figure electrodynamic tethered subsatellite experiments by the Space Shuttle. Vehicle charging due to dc beam emission up to 80 mA in the 150-200 km altitude range was repeatedly measured by both Langmuir and floating probes, and was found to be usually less than 10 V. During the 80 mA emission, clear evidence for the ignition of a beam-plasma discharge was obtained. In a tether deployment experiment, it was found that the tether wire acted as an antenna whose impedance decreased with the extension of the wire both in high-frequency and very low-frequency bands. Substantial rocket charging was observed during periods of electron-beam very low-frequency pulsing.

A method and apparatus for launching a spacecraft by use of a recoverable upper rocket stage

Abstract: An apparatus and method for launching a spacecraft including a payload and a delivery stage having a rocket engine powered by fluid bipropellant from the earth into a high energy orbit and for recovering the delivery stage are disclosed. By reducing the delivery stage mass, it becomes feasible and cost effective to recover the delivery stage for reuse. Delivery stage mass is reduced by several techniques including transporting the spacecraft and the fluid bipropellant to a parking orbit with the fluid bipropellant in tanks external to the spacecraft; transferring the fluid bipropellant to light weight tanks integral to the spacecraft; controlling the relative flow rates of the fluid bipropellant constituents to the rocket engine during firing of the rocket engine to ensure complete use of both bipropellant constituents; and controlling ascent and descent maneuvers from remote tracking stations. A space shuttle can be used to transport the spacecraft and fluid bipropellant in its cargo bay to the parking orbit and recover the delivery stage at the end of a mission. The invention is particularly useful for delivery of payloads to geosynchronous orbits.

Direct measurement of high frequency, solid propellant, pressure-coupled admittances

Abstract: This paper presents an experimental method that is capable of directly measuring solid propellant pressurecoupled responses at the high frequencies associated with tangential mode instabilities inside solid propellant rocket motors. The method utilizes a magnetic flowmeter to measure the velocity oscillation above a burning propellant surface simultaneously with a pressure oscillation measurement within an externally excited combustion chamber. A magnetic flowmeter burner was designed and constructed to evaluate this method of pressurecoupled response measurement. Response measurements were obtained for two formulations of AP/HTPB composite propellant at pressure oscillation frequencies of 4000 and 8000 Hz. The measurement data displayed repeatable trends in both the real and imaginary parts of the pressure-coupled response function.

Rocket drive with air intake

Abstract: A rocket propulsion system working with cryogenic fuel is modified by supersonic and hypersonic flight wherein air is sucked in to be used in lieu of the oxidizer in the rocket system. Hybrid operation is provided for.

Ocean launching apparatus of space rocket

Abstract: An ocean launching apparatus of space rockets comprises a launch pad platform having a floating-island structure and a semi-submersible hull for loading and transporting the launch pad platform. When the apparatus reaches an intended ocean area, a rocket set on the launch pad platform is floated from the hull in a semi-submersing state on the ocean together with the launch pad platform and moved to a launching site. In this way, it is possible to provide an ideal launching site of space rockets, increase an economical efficiency of launching of space rockets, and guarantee the safety if an accident should occur.

High solids ratio solid rocket motor propellant grains and method of construction thereof

Abstract: A high acceleration high performance solid rocket motor grain such as for a ballistic defense missile or rocket assisted projectile comprises a propellant material which includes a highly plasticized binder so that the grain has a solids ratio equal to at least about 95 percent. In order that the grain with such a solids ratio may have adequate strength and withstand high acceleration forces, a reticulated structure is embedded therein. A method of constructing a rocket motor having such a grain is also disclosed.

Evaluation of rocket plume signature uncertainties

Abstract: A method is developed with which to evaluate the uncertainty in predictions of the infrared signature of metalized fuel, solid-propellant rocket plumes. The method consists of 1) identifying parameters that represent the major sources of uncertainty, 2) evaluating the infrared signature as each parameter is varied about its nominal value, 3) determining the uncertainty interval of each parameter and 4) root sum squaring the uncertainties to obtain the uncertainty factor for the infrared signature. Uncertainties in the index of refraction, particle size, temperature and number density, gas mole fraction temperature, plume pressure and radius, and A12O3 melting temperature, are considered. These uncertainties are combined to yield an overall uncertainty in the plume infrared emission and a relative ranking of the importance of each uncertainty parameter. Numerical results are generated using the standardized infrared radiation model (SIRRM) code for a solid-propellant tactical rocket plume. The most important parameters in the tactical rocket signature uncertainty are 1) A12O3 particle temperature, the real part of the refractive index and plume radius at 90 deg aspect angle and 2) the real part of the Al2O3 particle refractive index, the A12O3 particle temperature and size, and plume size at a nose-on aspect angle.

A method and an apparatus for tracking a missile in its trajectory

Abstract: The disclosure relates to a method and an apparatus (2) for tracking an unguided rocket projectile during at least a part of its trajectory. According to the invention, it is possible, with the knowledge of the difference between the actual trajectory of the rocket projectile and its theoretically calculated trajectory, to decide on and execute those corrections which are required in order that subsequent rocket projectiles (3) launched at the same target area shall register hits. According to the invention, the rocket projectile (3) is tracked by means of an IR sensor, or alternatively, a TV camera (2) which senses the contrast between the projectile (3), or parts thereof, and preferably the nozzle of the rocket motor, and the background (the heavens).

Apollo Lightcraft Project

Abstract: The detailed design of a beam-powered transatmospheric vehicle, the Apollo Lightcraft, was selected as the project for the design course. The principal goal is to reduce the LEO payload delivery cost by at least three orders of magnitude below the Space Shuttle Orbiter in the post 2020 era. The completely reusable, single-stage-to-orbit shuttlecraft will take off and land vertically, and have a reentry heat shield integrated with its lower surface. At appropriate points along the launch trajectory, the combined cycle propulsion system will transition through three or four airbreathing modes, and finally use a pure rocket mode for orbital insertion. The objective for the Spring semester propulsion source was to design and perform a detailed theoretical analysis on an advanced combined-cycle engine suitable for the Apollo Lightcraft. The preliminary theoretical analysis of this combined-cycle engine is now completed, and the acceleration performance along representative orbital trajectories was simulated. The total round trip cost is $3430 or $686 per person. This represents a payload delivery cost of $3.11/lb, which is a factor of 1000 below the STS. The Apollo Lightcraft concept is now ready for a more detailed investigation during the Fall semester Transatmosphere Vehicle Design course.

Space station propulsion

Abstract: Two propulsion systems have been selected for the space station: gaseous H/O rockets for high thrust applications and the multipropellant resistojets for low thrust needs. These two thruster systems integrate very well with the fluid systems on the space station, utilizing waste fluids as their source of propellant. The H/O rocket will be fueled by electrolyzed water and the resistojets will use waste gases collected from the environmental control system and the various laboratories. The results are presented of experimental efforts with H/O and resistojet thrusters to determine their performance and life capability, as well as results of studies to determine the availability of water and waste gases.

Application of single crystal superalloys for earth-to-orbit propulsion systems

Abstract: Single crystal superalloys were first identified as potentially useful engineering materials for aircraft gas turbine engines in the mid-1960's. Although they were not introduced into service as turbine blades in commercial aircraft engines until the early 1980's, they have subsequently accumulated tens of millions of flight hours in revenue producing service. The space shuttle main engine (SSME) and potential advanced earth-to-orbit propulsion systems impose severe conditions on turbopump turbine blades which for some potential failure modes are more severe than in aircraft gas turbines. Research activities which are directed at evaluating the potential for single crystal superalloys for application as turbopump turbine blades in the SSME and advanced rocket engines are discussed. The mechanical properties of these alloys are summarized and the effects of hydrogen are noted. The use of high gradient directional solidification and hot isostatic pressing to improve fatigue properties is also addressed.

Small-Scale Features in the Middle Atmosphere Wind Field at Saskatoon, Canada (52°N, 107°W): An Analysis of MF Radar Data with Rocket Comparisons.

Abstract: The wind field of the middle atmosphere (60-100 kin) is sampled by the Saskatoon Medium Frequency radar: temporal resolution is normally 5 rain, and vertical is 1.5/3 kin. Profiles are analyzed for gravity waves (GW), and periods r from 10 rain-10 h are measured, with 3,z > 2 km and amplitudes > 5 m s-. The profiles are quite similar to those from rocket soundings. Wind vector shears are also consistent with "wind corners" evident in recent rocket data. Vertical shears of the horizontal wind and GW amplitudes (10 < r < 60 rain)are calculated and shown as annual height-time cross sections; values near 60 km are compared with rocket data from nearby Primrose Lake. Regions favoring dynamic and convective instability and GW saturation are located. The scattered radar power is shown as a seasonal cross section and compared to the shear and GW features. Finally, the dissipation rate of GW kinetic energy is calculated and compared with related MF radar and rocket wind estimations.

Performance of high-area-ratio nozzle for a small rocket thruster

Abstract: Theoretical estimates of supersonic nozzle performance have been compared to experimental test data for nozzles with an area ratio of 100:1 conical and 300:1 optimum contour, and 300:1 nozzles cut off at 200:1 and 100:1. These tests were done on a Hughes Aircraft Company 5 lbf monopropellant hydrazine thruster with chamber pressures ranging from 25 to 135 psia. The analytic method used is the conventional inviscid method of characteristic with correction for laminar boundary layer displacement and drag. Replacing the 100:1 conical nozzle with the 300:1 contoured nozzle resulted in an improvement in thrust performance of 0.74 percent at chamber pressure of 25 psia to 2.14 percent at chamber pressure of 135 psia. The data is significant because it is experimental verification that conventional nozzle design techniques are applicable even where the boundary layer is laminar and displaces as much as 35 percent of the flow at the nozzle exit plane.

Experiment on multiple fuel supplies to air breathing rocket combustors

Abstract: An experimental study on multiple fuel supplies to cylindrical subsonic mode combustors of air breathing rockets was made for the purpose of reducing combustor length. The experiment consisted of two parts, one in which all the fuel was supplied through a rocket with multiple nozzles and the other in which a portion of the fuel was directly fed into a secondary combustor with the rest being supplied through a rocket with a single nozzle. The results are compared with each other and with those of a reference experiment in which a rocket with a single nozzle without fuel injection was used. In the case with multiple nozzles, mixing and combustion efficiencies were higher than those in the reference experiment for the same combustor length. They collapse to a single curve against the combustor length nondimensionalized by the rocket nozzle exit diameter, except for the combustion efficiencies for unstable combustion. Increase of the injection mass flow rate makes mixing and combustion efficiencies rise, but it soon becomes less effective beyond a certain limit. Combustion instability observed in the reference case for long combustors is suppressed in both cases.

An analytical and experimental investigation of resistojet plumes

Abstract: As a part of the electrothermal propulsion plume research program at the NASA Lewis Research Center, efforts have been initiated to analytically and experimentally investigate the plumes of resistojet thrusters. The method of G.A. Simons for the prediction of rocket exhaust plumes is developed for the resistojet. Modifications are made to the source flow equations to account for the increased effects of the relatively large nozzle boundary layer. Additionally, preliminary mass flux measurements of a laboratory resistojet using CO2 propellant at 298 K have been obtained with a cryogenically cooled quartz crystal microbalance (QCM). There is qualitative agreement between analysis and experiment, at least in terms of the overall number density shape functions in the forward flux region.

Prediction of aluminum combustion efficiency in solid propellant rocket motors

Abstract: Aluminum combustion efficiencies in solid propellant rocket motors are calculated by means of an improved theoretical model for the physical and chemical processes involved. In particular, more complete allowance is made for the fact that the gas-phase concentration of oxidizing species in the rocket chamber varies from point to point as a result of depletion by the aluminum combustion reaction. Analyses are carried out for both onedimensional and two-dimensional gas flowfields and both with and without consideration of particle velocity lags. Numerical calculations indicate that the neglect of variations in the oxidizer concentration a leads to significant errors when the combustion efficiency is substantially less than unity. These errors can be effectively eliminated, with little additional effort, by accounting for the variations in a under the assumption that particles of different sizes may be treated independently and that oxidizer depletion is not seriously affected by velocity lags-

Consideration of two-dimensional flow effects on nozzleless rocket performance

Abstract: A NUMERICAL analysis procedure has been developed ./mfor examination of two-dimensional flow effects on nozzleless rocket motor performance. In this analysis, it is assumed that there are no radial variations in static pressure and that flow is isentropic along each streamtube. Performance predictions are compared with standard one-dimensional flow predictions. It is found that two-dimensional effects on performance are considerably less than predicted by previous analyses, with a maximum increase in vacuum throat specific impulse of only 2.0%. As expansion ratio considerations are brought in, the performance difference decreases to less than 1%.

Fluid mechanics of spinning rockets

Abstract: : This report presents the results of a detailed investigation of the influence of time-dependent combustion gas flows on the attitude dynamics of spinning rocket propelled space vehicles. The work was motivated by a need to understand the origins of a potentially serious system performance problem first detected in the PAM-D series of spin stabilized upper stages. Small wobbling (often referred to as nutation or coning ) induced during separation of the rocket motor burn. The growth ceased abruptly at motor burnout, and final cone angles as large as 17 deg were reached in some flights. The same phenomenon was encountered in two flights of the PAM-DII, a similar vehicle utilizing a larger motor. Conventional theories of spinning rocket dynamics failed to explain this behavior. Since the telemetry data shows that the severity of the problem depends on spacecraft mass properties and other system parameters, it is crucial that the origins of the instability be understood completely in order that serious mission degradation can be avoided in future orbit raising operations. A costly interim fix, which sidesteps the need to understand the physical origins of the problem, is the use of a strap-on nutation control system as used in the Air Force SGS II missions. This approach is in direct conflict with the philosophy of solid rocket space propulsion, which is based on its inherent simplicity and low cost.

Rocket/launcher interface

Abstract: Sabot adapters are provided for mounting along the outer periphery of a ret on opposite sides of the center of gravity with the sabot adapters at each location including inner and outer ring shaped structures that are made of segments and telescoped together with tapering surfaces so that as the two ring structures are moved together they expand radially to fill the radial spacing between the rocket and a rocket launch tube to provide support for the rocket and cause accurate firing of the rocket.

Method of and apparatus for casting solid propellant rocket motors

Abstract: A method of and apparatus for casting uncured solid propellant in solid propellant rocket motors includes a casting bayonet having annular slits formed in the exit end thereof that improve the fluidity and self-leveling behavior of the uncured propellant. Both of these effects eliminate or reduce the trapping of air in the cast propellant, thus providing a better product.

Flugzeug und verfahren zum lagern von brennstoffen in einem flugzeug

Abstract: A propulsion device includes a rocket casing 8 having a combustor chamber 64 with a combustor liner 52 and a nozzle throat chamber 62 with a nozzle throat liner 52; a fuel passageway 56 defined between the liners and an outerwall 54 and a plurality of vanes 80 disposed in the fuel passageway to direct fuel circumferentially whereby the fuel travels a longer path through the passageway, and the residence time of fuel in the passageway is increased to promote heat transfer to the fuel. A plurality of fuel injection holes 68 in the combustor liner are also arranged so that fuel is injected into the combustion chamber in a direction which promotes circumferential motion of the fuel. An oxidizer is introduced through oxidizer injection holes 66 at the center of the combustion chamber to inject oxidizer into the combustion chamber in a direction which promotes circumferential motion of the oxidizer. The fuel injection holes and oxidizer injection holes may be staggered to promote overlapping of fuel and oxidizer and further promote mixing prior to and during combustion. Details are given concerning the nature of the fuel and other propellant components, and catalyst assisted endothermic pyrolysis of the fuel in the passage way is referred to. Details are also given concerning the construction of the rocket casing walls and propellant storage tanks (Fig. 2) and the arrangement of tanks and rocket motors in a vehicle (Fig. 1). Use of the device in a ram-jet or scram-jet is also referred to.

Rocket-borne low energy ion measurements in space

Abstract: A capped hemisphere electrostatic analyzer has been developed for the purpose of performing detailed studies of charged particle distributions in space from sounding rocket platforms. This instrument employs micro channel plate detectors in conjunction with a linear resistive anode to carry out angular imaging, by resistive charge division, of particle arrivals. Two such instruments, capable of supplying 64 x 32 angle-energy positive ion distributions every $\sim1$ second were flown on two separate high latitude sounding rockets in February, 1985, from Sondre Stromfjord, Greenland. One of these two rockets featured an active ion beam experiment whereby 200 eV/q Ar$\sp{+}$ ions were injected into the ionospheric plasma from a separated sub payload in broad $(\sim60\sp\circ$ FWHM) beams directed alternately either parallel to or perpendicular to the geomagnetic field. Ion fluxes associated with beam operations were observed on the main payload out to a main/sub payload separation distance of nearly 1 km. Several distinct ion populations are identified, based on their energy/pitch angle characteristics and the existence of ion fluxes at unexpected energies and pitch angles is demonstrated and discussed in light of current understanding of these types of beam-plasma systems. The ion flux signatures of parallel versus perpendicular beam injections are compared and contrasted.