Space Shuttle Solid Rocket Motor Research Articles

Space Shuttle Solid Rocket Motor Plume Pressure and Heat Rate Measurements

The solid rocket booster main flame deflector at NASA Kennedy Space Center Launch Complex 39A was instrumented to measure heat rates, pressures, and temperatures on the final three space shuttle launches. Because the solid rocket booster plume is hot and erosive, a robust tungsten piston calorimeter was developed to compliment measurements made by off-the-shelf sensors. Witness materials were installed, and their melting and erosion response to the Mach 2, 4000°F, 4 s duration plume was observed. The data show that the specification used for the design of the main flame deflector thermal protection system overpredicts heat rates by a factor of three and underpredicts pressures by a factor of two. The discovery of short-duration heating spikes that occur when aluminum oxide slag solidifies on the main flame deflector explains this heat rate overprediction. This study allows improvement of solid rocket motor launch site and test stand computational fluid dynamics models and the concomitant slag deposition heat transfer models.

Development of Thermal Barriers for Solid Rocket Motor Nozzle Joints

The Space Shuttle solid rocket motor case assembly joints are sealed using conventional 0-ring seals. The 5500+F combustion gases are kept a safe distance away from the seals by thick layers of insulation. Special joint-fill compounds are used to fill the joints in the insulation to prevent a direct flowpath to the seals. On a number of occasions. NASA has observed in several of the rocket nozzle assembly joints hot gas penetration through defects in the joint- fill compound. The current nozzle-to-case joint design incorporates primary, secondary and wiper (inner-most) 0-rings and polysulfide joint-fill compound. In the current design, 1 out of 7 motors experience hot gas to the wiper 0-ring. Though the condition does not threaten motor safety, evidence of hot gas to the wiper 0-ring results in extensive reviews before resuming flight. NASA and solid rocket motor manufacturer Thiokol are working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design and a thermal barrier, This paper presents burn-resistance, temperature drop, flow and resiliency test results for several types of NASA braided carbon-fiber thermal barriers. Burn tests were performed to determine the time to burn through each of the thermal barriers when exposed to the flame of an oxy-acetylene torch (5500 F), representative of the 5500 F solid rocket motor combustion temperatures. Thermal barriers braided out of carbon fibers endured the flame for over 6 minutes, three times longer than solid rocket motor burn time. Tests were performed on two thermal barrier braid architectures, denoted Carbon-3 and Carbon-6, to measure the temperature drop across and along the barrier in a compressed state when subjected to the flame of an oxyacetylene torch. Carbon-3 and Carbon-6 thermal barriers were excellent insulators causing temperature drops through their diameter of up to a 2800 and 2560 F. respectively. Gas temperature 1/4 downstream of the thermal barrier were within the downstream Viton 0-ring temperature limit of 600 F. Carbon-6 performed extremely well in subscale rocket char motor tests when subjected to hot gas at 3200 F for an 11 second rocket firing, simulating the maximum downstream joint cavity fill time. The thermal barrier reduced the incoming hot gas temperature by 2200 F in an intentionally oversized gap defect, spread the incoming jet flow, and blocked hot slag, thereby offering protection to the downstream 0-rings.

Investigation of the acoustic properties of composite materials

An effort to provide inspection of the bondlines in the Space Shuttle Solid Rocket Motor (SRM) is discussed. Investigation of the bondlines in the membrane region of the SRM has centered on representative samples provided by Morton Thiokol. These samples have pull-tabs, or in one case grafoil inserts, to simulate delaminations at various points in the multilayer structure of steel case, liner, insulation, liner, and propellant. Discussed here is the continuing approach to examining these defects which is predicated on a model of an ultrasonic wave propagating in a lossy multilayer resonant structure, the experimental arrangements used, and the results of these experiments.

Redesigned solid rocket motor case-to-nozzle joint development

The case-to-nozzle joint in the Space Shuttle solid rocket motor is designed to provide thermal protection to the case and nozzle metal parts and to the O-ring seals during motor operation and subsequent heat soak. As a part of the redesign effort, several design improvements were incorporated into the joint configuration. These included an unvented bonded joint, the addition of a pressure-actuated flap gap next to the bondline, the addition of an adhesive wiping O-ring, and the addition of 100 bolts in the radial direction to restrict joint movement. With the new design features, it became necessary to develop a new nozzle installation procedure. Also, a number of hot-fire static tests were performed before the joint could be used in a flight motor. In several of these tests, flaws were deliberately placed in the case-to-nozzle bondline and O-rings to determine how the redesigned joint would function with a portion of the redundant sealing system disabled prior to motor operation. Each flaw condition tested was accompanied by a flow/thermal analysis. Once these analyses had been verified, analyses were then used to examine other flaw configurations that were not being tested.

Calculation of the two-phase aft-dome flowfield in solid rocket motors

The two-phase flowfield in the aft-dome region of a solid rocket motor (SRM) with submerged nozzle has been simulated using a combined Eulerian-Lagrangian analysis. This analysis uses the numerical solution of ensemble averaged Navier-Stokes equations for the continuous (gas) phase coupled with a Lagrangian analysis for the discrete (particulate) phase to simulate the two-phase internal flow. A linearized block-implicit (LBI) scheme is used to solve the governing equations for the continuous phase, which allows the use of a highly stretched grid with sublayer resolution. The motion of the particles is tracked in computational coordinate space resulting in computational efficiency, and the interphase coupling terms for the Eulerian analysis are computed from the instantaneous distribution of the particles. A low Reynolds number form of the k-e turbulence model is used with modifications for injection driven flows. Calculations have been performed for a particular grain configuration of the Space Shuttle SRM. The flowfield in the vicinity of the submerged nozzle, the particle trajectories, and the sensitivity of two-phase effects (such as slag accumulation) to the particle injection parameters are presented in this paper.

Effect of idealized asymmetric inhibitor stubs on circumferential flow in the Space Shuttle SRM

Computational fluid dynamic analyses have been performed to calculate circumferential pressure and velocity gradients in the vicinity of an asymmetric inhibitor stub in the port of the Space Shuttle solid rocket motor. The three-dimensional Navier-Stokes equations were solved by an iterative finite volume algorithm, SIMPLE, incorporated in a general purpose computer code, FLUENT. Turbulence was represented by way of effective viscosity, which was calculated from local turbulence energy and its dissipation rate (K-e model). The numerical predictions were compared with the measurements from a 1.5% scale, cold-flow model of the redesigned solid rocket motor. The calculated circumferential pressure distribution upstream of the inhibitor stub compares well with the measured data.

Space Shuttle solid rocket motor aft-end internal flows

Flow-visualization techniques were employed in a Vs-scale model of the Space Shuttle solid rocket motor to investigate the dominant aft-end internal flow patterns. The model included the two aft segments, the aft dome, and the convergent portion of the gimballed nozzle. The effects of blowing, which would result from gases produced by the burning propellant, were simulated through the introduction of a uniform distribution of water along simulated burn-back patterns representing the surface at three different times in a firing. It was found that the effects of vortices, shed from protruding inhibitor sections, upon flow a few diameters downstream were greatly diminished by the effects of wall injection. It was also found that the extent of circumferential flow resulting from the removal of a portion of protruding inhibitor was limited to the region near the missing portion. Strong circumferential flow in the aft dome was observed when no grain surface was present in the dome. This flow included nozzle-entrance vortices, which appear to result from interactions between the circumferential and toroidal flows associated with the submerged nozzle and the aft dome.

Navier-Stokes analysis of solid propellant rocket motor internal flows

A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solid propellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.

Modeling of the Space Shuttle solid rocket motor nozzle boot cavity pressurization process

Pressurization of the Space Shuttle solid rocket motor (SRM) nozzle boot cavity involves compressible gas flow along a duct with friction and a pressurizing volume with heat transfer at the boundaries. Numerical simulation of the process requires a solution of the simultaneous compressible flow, boot cavity conservation, and solid-conduction equations. Basic geometry consists of circular flow path connected to a plenum-free volume that is surrounded by a solid-conduction media (i.e., the cowl vent holes, boot cavity, and cavity insulation). An existing methodology is used to compute the compressible mass-flow rates through the vent hole. Boot cavity conditions are computed by the application of a finite-difference form of the conservation equations. In the boot cavity insulation (i.e., the solid-conduction regions), the finite-difference transient diffusion equation is applied. A conjugate solution is developed whereby the coupled first-order differential equations are solved simultaneously, along with a set of algebraic equations, by fully implicit successive point iteration. A model was constructed to simulate a boot cavity pressurization test, and a comparison of model response vs test data is given. The agreement is generally good.

Ignition/duct overpressure induced by Space Shuttle solid rocket motor ignition

An analytical methodology, with a propagating solid rocket motor (SRM) exhaust front as a perturbation pressure wave generator, is developed to enhance understanding of the ignition/duct overpressure (IOP/DOP) induced during the Space Shuttle liftoff. Waveform, amplitude, and low-frequency responses of IOP/DOP are simulated. For a full-scale configuration, IOP and DOP responses are clearly isolated and distinguishable. However, in the subscale tests, an interaction region of IOP and DOP exists, which obscures interpretation of the test data. The methodology offers a first-order assessment of these phenomena with respect to the launch complex geometry and the SRM ignition transient, and provides guidance for the test data interpretation.

Thrust and ignition transients of the Space Shuttle solid rocket motor

Techniques developed for predicting and analyzing ignition, pressurization, and thrust transients in large segmented motors were applied to the four developmental Space Shuttle solid rocket motors. Following the first test, attention was focused on understanding dynamic thrust and the rate of thrust increase during pressurization and predicting effects of reduced igniter mass flow rate and gas temperature. The close coupling between the igniter and the head-end segment and the high length-to-diameter ratio make understanding of the longitudinal pressure waves essential. The igniter performance along with igniter to head-end segment geometry govern the induction interval and head-end segment ignition. Following ignition of the head-end segment, subsequent flame spreading and pressurization tend to be motor properties which are not appreciably altered by changes in the head-end igniter mass flow rate.

Effects of Propellant Temperature Gradients on Thrust Imbalance of the Space Shuttle

The internal ballistic effects of combined radial and circumferential grain temperature gradients are evaluated theoretically for the Space Shuttle solid rocket motors (SRM's). A simplified approach is devised for representing with closed-form mathematical expressions the temperature distribution resulting from the anticipated thermal history prior to launch. The internal ballistic effects of the gradients are established by use of a mathematical model which permits the propellant burning rate to vary circumferentially. Comparative results are presented for uniform and axisymmetric temperature distributions and the anticipated gradients based on an earlier two-dimensional analysis of the center SRM segment. The thrust imbalance potential of the booster stage is assessed based on the difference in the thermal loading of the individual SRM's of the motor pair which may be encountered in both summer and winter environments at the launch site. Results indicate that grain temperature gradients could cause the thrust imbalance to be approximately 10% higher in the Space Shuttle than the imbalance caused by SRM manufacturing and propellant physical property variability alone.