Showing papers on "Liquid-propellant rocket published in 1975"

Solid rocket motor nozzles

Abstract: The steps in the nozzle design process are examined. The nozzle designer's role in defining design requirements and constraints is included along with discussions of each of the three basic phases of the nozzle design process itself: (1) aerodynamic design, in which the gas-contacting surfaces are configured to produce the required performance within the envelope limits; (2) thermal design, in which termal liners and thermal insulators are selected and configured to maintain the surfaces as closely as practical against effects of erosion and to limit the structure temperature to acceptable levels; and (3) structural design, in which materials are selected and configured to support the thermal components and to sustain the predicted loads. Analytical techniques that are used to establish thermal and structural design integrity and to predict nozzle performance are discussed along with methods for nozzle quality assurance. Emphasis is placed on nozzle design and materials for modern high-temperature aluminized propellants. Recurring nozzle design problems of graphite cracking and ejection, differential erosion at material interfaces, lack of sufficient proven nondestructive testing (NDT) techniques, the uncertainty of adhesive bonding, and inadequate definition of material properties, particularly at high temperatures are considered.

Scaling of Rocket Nozzle Admittances

Abstract: To apply the admittance data measured experimentally for small-scale nozzles, the necessary scaling criteria must be established. The results of an investigation undertaken to determine these criteria are presented and the data indicate that under cold-flow conditions the damping of axial instabilities provided by full-scale nozzles can be determined from the attenuation data obtained experimentally for geometrically similar small-scale nozzles. The scaling of both short and long nozzles has been investigated. The experimental data also support the scaling criteria proposed by Crocco in his analytical investigation of nozzle damping.

Solid Rocket Motor for the Space Shuttle Booster

Abstract: The Solid Rocket Motor (SRM) for the Space Shuttle Solid Rocket Booster (SRB) is being developed by Thiokol Corporation, Wasatch Division under the cognizance of Marshall Space Flight Center (MSFC), Contract NAS8-30490. Solid Rocket Motor requirements are summarized, and the overall SRM configuration and performance characteristics are presented in this paper. Design details of the major SRM components are reviewed and the approach taken in each component area is identified and related to previous industry experience. The overall scope of the SRM Project is presented. Activities on the project since authorization to proceed in June 1974 are reviewed. A summary schedule is presented which relates key SRM Project activities to the overall Shuttle Program.

Solid rocket technology advancement for Space Tug and IUS applications

Abstract: Two-burn restartable solid propellant rocket motors for the kick stage (auxiliary stage) of the Shuttle Tug, or Interim Upper Stage, are described, with details on features and test results of the ignition and quench (thrust termination) systems and procedures, fabrication of propellant and insulation, explosion hazards of propellants, and comparative data on present and future motor design. These rocket motor systems are designed for upper stage augmentation of launch vehicles and possible service in Shuttle-launched outer planet spacecraft.

Study of liquid oxygen/liquid hydrogen auxiliary propulsion systems for the space tug

Abstract: Design concepts are considered that permit use of a liquid-liquid (as opposed to gas-gas) oxygen/hydrogen thrust chamber for attitude control and auxiliary propulsion thrusters on the space tug. The best of the auxiliary propulsion system concepts are defined and their principal characteristics, including cost as well as operational capabilities, are established. Design requirements for each of the major components of the systems, including thrusters, are developed at the conceptual level. The competitive concepts considered use both dedicated (separate tanks) and integrated (propellant from main propulsion tanks) propellant supply. The integrated concept is selected as best for the space tug after comparative evaluation against both cryogenic and storable propellant dedicated systems. A preliminary design of the selected system is established and recommendations for supporting research and technology to further the concept are presented.

Thrust vector control by liquid injection for solid propellant rockets

Abstract: In liquid injection thrust vector control, a rocket jet is deflected for steering purposed by injecting a liquid into the nozzle exit cone. The liquid is preferably both dense and reactive so that it adds mass and energy and generates shocks in the supersonic exhaust. This behavior increases thrust in the affected part of the jet producing not only a side force for steering but an addition to axial thrust. This paper presents a summary of current liquid injection thrust vector control technology, including procedures for design, development, analysis, testing and evaluation, together with supporting data and references.

Space shuttle Solid Rocket Booster (SRB) separation

Abstract: The system is described which is used to separate the solid rocket boosters from the space shuttle after they have expended most of their propellant and their thrust is near burnout. The dynamics of the separation are simulated in a computer program so that the separation system can be analyzed. The assumptions and ground rules used in analyzing this system are explained and the method of analysis is delineated. The capability of the separation system is presented together with data which may be used to aid in the design of the external tank and solid rocket booster interface. The results of a parameter study to determine the sensitivity of the separation to the initial state of the space shuttle are also presented.

Reverse flow aft inlet ramjet system

Abstract: A combined rocket/ramjet for propelling a missile including a missile payload module and an integral rocket/ramjet engine system strapped on the missile payload module, and incorporating a liquid fuel rocket engine and a ramjet engine having a combustion chamber common with the rocket engine, an exhaust nozzle, and multiple aft-mounted air inlets movable between a retracted position during rocket boost flight and an extended, pop-out position to receive and direct ram air in a reverse flow direction into the combustion chamber during ramjet flight.

Thrust vector control for the Space Shuttle Solid Rocket Motor

Abstract: Thrust vector control (TVC) for the Space Shuttle Solid Rocket Motor (SRM) is obtained by omniaxis vectoring of the nozzle The development and integration of the system are under the cognizance of Marshall Space Flight Center (MSFC) The nozzle and flexible bearing have been designed and will be built by Thiokol Corporation/Wasatch Division The vector requirements of the system, the impact of multiple reuse on the components, and the unique problems associated with a large flexible bearing are discussed The design details of each of the major TVC subcomponents are delineated The subscale bearing development program and the overall development schedule also are presented

Effect of Nozzle Nonlinearities upon Nonlinear Stability of Liquid Propellant Rocket Motors

Abstract: A three dimensional, nonlinear nozzle admittance relation is developed by solving the wave equation describing finite amplitude oscillatory flow inside the subsonic portion of a choked, slowly convergent axisymmetric nozzle. This nonlinear nozzle admittance relation is then used as a boundary condition in the analysis of nonlinear combustion instability in a cylindrical liquid rocket combustor. In both nozzle and chamber analyses solutions are obtained using the Galerkin method with a series expansion consisting of the first tangential, second tangential, and first radial modes. Using Crocco's time lag model to describe the distributed unsteady combustion process, combustion instability calculations are presented for different values of the following parameters: (1) time lag, (2) interaction index, (3) steady-state Mach number at the nozzle entrance, and (4) chamber length-to-diameter ratio. In each case, limit cycle pressure amplitudes and waveforms are shown for both linear and nonlinear nozzle admittance conditions. These results show that when the amplitudes of the second tangential and first radial modes are considerably smaller than the amplitude of the first tangential mode the inclusion of nozzle nonlinearities has no significant effect on the limiting amplitude and pressure waveforms.

Linear rocket engines for advanced space transportation systems. [post-shuttle single-stage to orbit vehicles

Abstract: The linear rocket engine is shown to be a viable candidate propulsion system for post-Space Shuttle single-stage-to-orbit systems. The linear engine system has been developed and fired demonstrating high performance and long life with firing durations exceeding 500 seconds. The application of the split or dual combustor to the linear engine permits the uses of two different propellant combinations in a single engine system. The split combustor possesses the advantages of the two position extendible bell nozzle in a fixed nozzle configuration. Engine power cycles and applications to typical vehicles are discussed.