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Plans for conducting Phase C/D for a solid rocket motor booster vehicle are presented. Methods for conducting this program with details of scheduling, testing, and program management and control are included. The requirements of the space shuttle program to deliver a minimum cost/maximum reliability booster vehicle are examined.
The design, development, production, and launch support analysis for determining the solid propellant rocket engine to be used with the space shuttle are discussed. Specific program objectives considered were: (1) definition of engine designs to satisfy the performance and configuration requirements of the various vehicle/booster concepts, (2) definition of requirements to produce booster stages at rates of 60, 40, 20, and 10 launches per year in a man-rated system, and (3) estimation of costs for the defined SRM booster stages.
Results of a study to demonstrate the applicability of boron-epoxy-composite-reinforced titanium tubular members to a space shuttle booster thrust structure are presented and discussed. The experimental results include local buckling of all-composite and composite-reinforced-metal cylinders with low values of diameter-thickness ratio, static tests on composite-to-metal bonded step joints, and a test to failure of a boron-epoxy-reinforced titanium demonstration truss. The demonstration truss failed at 118 percent of design ultimate load. Test results and analysis for all specimens and the truss are compared. Comparing an all-titanium design and a boron-epoxy-reinforced-titanium (75 percent B-E and 25 percent Ti) design for application to the space shuttle booster thrust structure indicates that the latter would weigh approximately 24 percent less. Experimental data on the local buckling strength of cylinders with a diameter-thickness ratio of approximately 50 are needed to insure that undue conservatism is not used in future designs.
The study of solid rocket motors for a space shuttle booster was directed toward definition of a parallel-burn shuttle booster using two 156-in.-dia solid rocket motors. The study effort was organized into the following major task areas: system studies, preliminary design, program planning, and program costing.
An experimental wind tunnel investigation on the Marshall Space Flight Center's (MSFC) booster configuration of the space shuttle system is discussed. Tests were conducted in the MSFC 14" x 14" trisonic wind tunnel. Parameters investigated include; (1) a landing study at Mach number of .3 with the booster at three discrete heights above a ground plane (maximum height above ground plane was six inches) and: (2) a static stability study at subsonic Mach numbers, which included eleven deflections of 0, 10, 20 and 30 degrees with canard deflections of 0, 5, 10 and 15 degrees. Both a trapezoidial and a delta shaped canard were investigated at the above conditions; a grit size study was made over the entire Mach range from .3 to 5.0.
Selection of the most appropriate thermal protection system design approaches requires evaluation of alternate concepts. Such an evaluation was conducted for a representative booster and orbiter configurations. TPS concept examined for the booster included heat sink, metallic reradiative, reusable surface insulation, and surface cooled types. For the orbiter, emphasis was placed on metallic reradiative and reusable surface insulation concepts with uncooled, phase change material cooled, and water-cooled substructures. Three insulation materials of different densities were examined for the orbiter reradiative TPS to identify the weight sensitivity of various substructural design approaches to this variable.
Experimental aerodynamic investigations have been made on a .0035 scale model North American Rockwell/General Dynamics version of the space shuttle in the NASA/MSFC 14 x 14 Inch Trisonic Wind Tunnel. Static stability and control data were obtained on the delta wing booster alone (B-20) and with the delta wing orbiter (134D) mounted in various positions on the booster. Six component aerodynamic force and moment data were recorded over an angle of attack range from -10 to 24 deg at 0 and 6 deg sideslip angles and from -10 to +10 deg sideslip at 0 deg angle of attack. Mach number ranged from 0.6 to 4.96.
Mass properties data are presented to describe the characteristics of an expendable second stage with a reusable space shuttle booster. The final mass characteristics of the vehicle configurations for three specified payloads are presented in terms of weight, center of gravity, and mass moments of inertia. Three basic subjects are the integrated vehicle system, the expendable second stage, and the booster modifications.
Contract end item specifications required to define hardware end items for the expendable second stage on a reusable shuttle booster system are provided. The specifications are as follows: (1) expendable second stage, (2) space shuttle booster, and (3) space shuttle ground support system.
The general and design criteria for the expendable second stage of the reusable space shuttle booster are presented. The subjects discussed are: (1) trajectory development, (2) aerodynamic heating on basic payloads, (3) vehicle configurations, trajectory, and heat regimes, (4) orbit maneuvering system, (5) structural analysis, and (6) booster loads and structural sizing.
The results of analytical and experimental parametric studies of space shuttle booster, orbiter and launch vehicle aerodynamics are described. During this study over 1700 hours of experimental wind tunnel tests were conducted on several versions of the shuttle booster, orbiter and launch vehicle. Fifteen separate tests were conducted in three different test facilities. Due to the number of test programs conducted and the time required for test preparation, analysis of the test data has been limited to that required to drive the experimental program. A brief description of each of the experimental tests conducted including the test purpose and approach is included. Several test models were designed and fabricated in support of the experimental program. These models are described.
The evolution of the space shuttle booster system is reviewed from its initial concepts based on liquid-propellant reusable boosters to the final selection of recoverable, solid-fuel rocket motors. The rationale associated with each of the several major decisions in the evolution process is discussed. It is shown that the external tank orbiter configuration emerging from the latest studies takes maximum advantage of the solid rocket motor development experience and promises to be the optimum configuration for fulfilling the paramount shuttle program requirements of minimum total development risk within acceptable costs.
The performance requirements, preliminary designs, and development program plans for an airborne recovery system for the space shuttle solid rocket booster are discussed. The analyses performed during the study phase of the program are presented. The basic considerations which established the system configuration are defined. A Monte Carlo statistical technique using random sampling of the probability distribution for the critical water impact parameters was used to determine the failure probability of each solid rocket booster component as functions of impact velocity and component strength capability.
The results are presented of a space shuttle plume impingement study for the Manned Spacecraft Center configuration. This study was conducted as two tasks which were to (1) define the orbiter main stage engine exhaust plume flow field, and (2) define the plume impingement heating, force and resulting moment environments on the booster during the staging maneuver. To adequately define these environments during the staging maneuver and allow for deviation from the nominal separation trajectory, a multitude of relative orbiter/booster positions are analyzed which map the region that contains the separation trajectories. The data presented can be used to determine a separation trajectory which will result in acceptable impingement heating rates, forces, and the resulting moments. The data, presented in graphical form, include the effect of roll, pitch and yaw maneuvers for the booster. Quasi-steady state analysis methods were used with the orbiter engine operating at full thrust. To obtain partial thrust results, simple ratio equations are presented.
Several technical areas were encompassed in providing support for booster thermal environment test work. These areas included: (1) cavity flow heating, (2) rarefied flow heating, and (3) impulse operated model research and testing. Cavity flow heating problems were studied with respect to the proposed altitude control motors for the space shuttle. Available literature on this subject was reviewed and analytical predictive methods were summarized for use in planning testing work. Rarefied flow heating data was reviewed and correlated. The study showed the importance of considering rarefied flow conditions in launch thermal environment prediction. Impulse operated model research and testing was conducted to provide a basis for understanding and designing such models for booster thermal environment testing.
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.
The separation characteristics of the space shuttle solid rocket boosters (SRBs) are introduced along with the system level requirements for the booster separation motors (BSMs). These system requirements are then translated into specific motor requirements that control the design of the BSM. Each motor component is discussed including its geometry, material selection, and fabrication process. Also discussed is the propellant selection, grain design, and performance capabilities of the motor. The upcoming test program to develop and qualify the motor is outlined.
A model is described which provides a means for analyzing the complexities of ignition transients and pressure peaks of large, high performance, segmented solid rocket boosters. The method accounts for: (1) temporal and spatial development of the flow field set up by the head end igniter discharge, (2) ignition and flame spreading coupled to chamber flow, (3) the steep velocity, pressure, and temperature gradients that occur during the early phases of ignition, and (4) the interactions that produce ignition spikes (i.e., compression of chamber gases during pressurization, erosive burning, and mass added effect of igniter discharge). The technique differs from earlier models in that the flow interactions between the slots and main chamber are accounted for, and the original computer program for monolithic motors is improved. The procedures were used to predict the ignition transients of the current design for the space shuttle booster.