V. R. Sanal Kumar

Flame Spread with Sudden Expansions of Ports of Solid Propellant Rockets

Detailed theoretical and experimental studies on flame spread over nonuniform ports of solid propellant rockets have been carried out. An idealized two-dimensional laboratory motor was used for the experimental study with the aid of cinematography. A detailed numerical simulation of the flame spread has also been carried out with the help of a two-dimensional Navier-Stokes solver. Experimental results showing the phenomenon of secondary ignition have been reported earlier and also reviewed here with the inclusion of additional results of a three-dimensional geometry closer to a dual-thrust motor. In this paper more tangible results including the numerical modeling of flame spread have been reported. It has been shown conclusively that under certain conditions of step location, step height, and port height, which govern the velocity of gases at the step by the partially ignited propellant surface or by the igniter gas Row, secondary ignition can occur far downstream of the step, This is very likely to be within the recirculating Row region. The secondary ignition gives rise to two additional flame fronts, one of which spreads backward st a relatively lower velocity, presumably as a result of low reverse velocities present in the separation zone. This phenomenon is Likely to play an important role in the starting transient of solid propellant rockets with nonuniform ports.

Thermoviscoelastic characterization of a composite solid propellant using tubular test

The accurate thermoviscoelastic characterization of a solid propellant is critical for optimum design of grains for high-performance solid motors for space applications. A more realistic experimental method is reported for the characterization of propellants at multiaxial stress conditions using a grain-pressurization test called the tubular test. An idealized cylindrical grain with hydroxyl-terminated-polybutadiene-based propellant is used. From the measured outer radial expansion of the tubular specimen, creep compliance and the relaxation modulus of the propellant are computed based on linear viscoelastic theory. The relaxation-modulus data obtained from the tubular test are shown to be higher than the traditional uniaxial and strip biaxial tests at the same strain level. Different isothermal tubular tests are carried out at various strain levels, and master curves are generated. The tubular test with internal pressurization is found to be an easier, more inexpensive, and realistic method for predicting the relaxation-modulus values of solid propellants and to have great potential for the thermoviscoelastic characterization of solid propellants.

Starting Transient Flow Phenomena in Inert Simulators of Solid Rocket Motors with Divergent Ports

The basic idea behind a solid rocket motor (SRM) is simple but its design is a complex technological problem requiring expertise in diverse subdisciplines to address all of the physics involved. The design optimization of high-performance rockets is more complex when the mission demands dual thrust. The motivation for the present study emanates from the desire to explain the phenomena or mechanism(s) responsible for the high ignition peak pressure (pressure peak), pressure-rise rate, instabilities, and pressure oscillations often observed during the static tests and the actual flights of certain class of high-performance SRMs with nonuniform ports [1–9]. In the SRM industry many dual-thrust motors (DTMs) are known to have experienced abnormal high ignition peak pressure often on the order of 5 times the steady state value [6]. Various measures were taken to eliminate the peak pressure, but none of the conventional remedies seemed to help. Nevertheless, through the empirical techniques increasing the port area of the motor has been proposed as one of the remedies for reducing the unusual ignition peak of the DTM. Although such a remedy could negate the unacceptable peak pressure, it has affected the high-performance nature of the motor. Hence the elimination of the unusual ignition peak and the pressure-rise rate without sacrificing the basic grain configuration or the volume loading became a meaningful objective for further studies.

Studies on Internal Flow Choking in Dual-Thrust Motors

Adetailed picture of the internal flow during the starting transient of high-performance solid rocket motors (SRMs) is of topical interest for several reasons in addition to the motor performance itself [1–12]. Despite the fact that many of the existing models could predict the internal flow features of certain classes of SRMs, none of these models could capture the unusual starting transient flow features such as pressure overshoot and pressure-rise rate often observed during the initial phase of operation of the dual-thrust motors (DTMs) [1]. Ikawa and Laspesa [8] reported that during the first launching of the space shuttle from the Eastern Test Range, the launch vehicle experienced the propagations of a strongly impulsive compression wave. This wave was induced by the SRM ignition and was emanating from the large SRM duct openings. The analysis further showed that the compression wave created by ignition of the main grain was the cause of the ignition overpressure on the launch pad [9]. Alestra et al. [10] reported that Ariene 5 launcher experienced overpressure load during the liftoff phase. The overpressure is composed of the ignition overpressure, which emanates from the launch pad, and the duct overpressure, which emanates from the launch ducts. Of late, Sanal Kumar et al. [1,2] reported that abnormal high-pressure overshoot in certain class of DTMs during the startup transient is due to the formation of shock waves because of the fluid-throat effect, which has received considerable attention in the scientific community. This manuscript is the continuation of the previous connected note for establishing the intrinsic flow physics pertinent to internal flow choking in inert simulators of dual-thrust motors [1]. Note that the illustration of ignition pressure spike is deliberately set aside in this note for explaining the intrinsic flow physics pertinent to internal flow choking without complications arising from the propellant combustion.

Boundary-Layer Effects on Internal Flow Choking in Dual-Thrust Solid Rocket Motors

Theoretical studies have been carried out to examine internal flow choking in the inert simulators of a dual-thrust motor. Using a two-dimensional k-omega turbulence model, detailed parametric studies have been carried out to examine aerodynamic choking and the existence of a fluid throat at the transition region during the startup transient of dual-thrust motors. This code solves standard k-omega turbulence equations with shear flow corrections using a coupled second-order-implicit unsteady formulation. In the numerical study, a fully implicit finite volume scheme of the compressible, Reynolds-averaged, Navier-Stokes equations is employed. It was observed that, at the subsonic inflow conditions, there is a possibility of the occurrence of internal flow choking in dual-thrust motors due to the formation of a fluid throat at the beginning of the transition region induced by area blockage caused by boundary-layer-displacement thickness. It has been observed that a 55% increase in the upstream port area of the dual-thrust motor contributes to a 25% reduction in blockage factor at the transition region, which could negate the internal How choking and supplement with an early choking of the dual-thrust motor nozzle. If the height of the upstream port relative to the motor length is too small, the developing boundary layers from either side of the port can interact, leading to a choked,flow. On the other hand, if the developing boundary layers are far enough apart, then choking does not occur. The blockage factor is greater in magnitude for the choked case than for the unchoked case. More tangible explanations are presented in this paper for the boundary-layer blockage and the internal flow choking in dual-thrust motors, which hitherto has been unexplored.

Studies on Ejector Systems for Hydrogen Fuel Cell

An ejector is a pumping device which exchanges energy between a high energy primary fluid and a relatively low energy secondary fluid to produce a discharge of intermediate specific energy level but higher mass flow rate. The present study addresses a new method to control the performance of a variable ejector system which is applied to a hydrogen fuel-cell system to economize the fuel consumption. Although the variable ejector concept had been proved, its flow characteristics have not yet been established for meeting the industrial demands. Towards achieving this objective, in this paper, detailed experimental and numerical studies have been carried out and the characteristic curves are generated to understand the effects of the ejector throat area ratio and the operating pressure ratio on the entrainment of secondary stream. In the experimentation a movable cylinder, inserted into a conventional ejector-diffuser system, is used to change the ejector throat area ratio, which controls the mass flow rate of the suction flow. In the numerical study, a fully implicit finite volume scheme of the compressible, Reynolds-Averaged, Navier-Stokes equations employed. The results show that the variable ejector can control the recirculation ratio by changing the throat area ratio and the operating pressure ratio.

Simulation of Flame Spread and Turbulent Separated Flows in Solid Rockets

Theoretical studies have been carried out to examine the flame spread and the turbulent separated flows in solid rocket motors with non-uniform port geometry. Detailed parametric studies have been carried out to examine the influence of port geometry on flow separation and reattachment using a standard k-omega turbulence model. In solid rockets, the flow separation and the reattachment will cause secondary ignition followed by multiple flame fronts. An error in pinpointing the location of secondary ignition can lead to significant errors in the thrust transient prediction of solid rockets. The present study is expected to aid the designer for conceiving the physical insight on secondary ignition into problems associated with the prediction and the reduction of the peak pressure and the pressurization rate during the starting transient period of solid rocket motors with non-uniform ports.

Diagnostic Investigation of Instabilities and Pressure Oscillations in High-Performance Solid Rockets

Using a standard k-ω turbulence model, in this paper detailed numerical computations have been carried out in inert simulators to examine the geometric dependence of transient flow features of solid rocket motors. We observed that the damping of the temperature fluctuation is faster in solid rocket with convergent port than with divergent port geometry. We inferred that an early damping of the transient flow fluctuations using the port geometry is a meaningful objective for the suppression and control of the instability and/or pressure/thrust oscillations during the staring transient of solid rockets.

Ratiocinative Approach to Ignition Transient Modeling in Solid Rockets

Detailed theoretical studies have been carried out to examine the variation of ignition transient on identical motors with different ignition delay. Igniter ballistics and propellant properties are varied for getting different ignition delay. We observed that the altered variation of ignition delay will alter the flame spread and thereby the overall ignition transient history of identical solid rockets. Significant peak pressure variation is observed in the case of solid motors with non-uniform ports. We concluded that, after maintaining the constant igniter ballistics, the designer can reduce the peak pressure by altering the surface condition by pasting high conducting material over the surface of the propellant. Such a change in thermal conductivity can be easily achieved by slightly altering the metal loading in the propellant. This ratiocinative approach will help the designer in reducing the peak pressure by altering the propellant ignition properties and the igniter characteristics without sacrificing the basic grain configuration or the volume loading.

A Computational Characterization of the Supersonic Coherent Jet

§¶ Supersonic coherent jet represents a break-through technology that offers many industrial applications. This paper provides an overview of conventional supersonic jets and reviews fundamental aspects of coherent jets. The present study also investigates the major characteristics of the supersonic coherent jet. A computational study is carried out using a two dimensional Navier-Stokes solver with standard k-H turbulence model. Three different kinds of supersonic jets, which are a conventional supersonic jet, a supersonic coherent jet and a supersonic coflow jet, are investigated for the purpose of comparison of the major characteristics of supersonic jet. We observed through the numerical studies that the jet core length in a supersonic coherent jet is 1.8 times longer than in the conventional supersonic jet. The width and half-width of the jet is the smallest in supersonic coherent jet. We concluded that the entrainment effect in supersonic coherent jet is the smallest while comparing with the conventional supersonic jet and supersonic coflow jet.