Martin J. Chiaverini

Regression Rate Behavior of Hybrid Rocket Solid Fuels

An experimental investigation of the regression-rate characteristics of hydroxyl-terminated polybutadiene (HTPB) solid fuel burning with oxygen was conducted using a windowed, slab-geometry hybrid rocket motor. A real-time, x-ray radiography system was used to obtain instantaneous solid-fuel regression rate data at many axial locations. Fuel temperature measurements were made using an array of 25- πm e ne-wire embedded thermocouples. The regression rates displayed a strong dependence on axial location near the motor head-end. At lower mass e ux levels, thermal radiation was found to signie cantly ine uence the regression rates. The regression rates werealso affected by theadditionofactivated aluminum powder.A 20%by weightaddition of activated aluminum to HTPB increased the fuel mass e ux by 70% over that of pure HTPB. Correlations were developed to relate the regression rate to operating conditions and port geometry for both pure HTPB and for HTPB loaded with certain fractions of activated aluminum. Thermocouple measurements indicated that the fuel surface temperatures for pureHTPBwerebetween930 and1190 K.TheHTPBactivationenergywasestimatedat11.5 kcal/mole,suggesting that the overall regression process is governed by physical desorption of high-molecular weight fragments from the fuel surface.

Solid-Fuel Regression Rate Behavior of Vortex Hybrid Rocket Engines

Aseriesofstaticenginee ringswereconductedtoinvestigatethesolid-fuelregressionratebehaviorandoperating characteristicsofvortexhybridrocketengines.Thevortexhybridenginecone gurationischaracterizedbyacoaxial, coswirling, countere owing vortex combustion e eld in a cylindrical fuel port. To generate this e owe eld, oxidizer is injected through a swirl injector located between the aft end of the fuel grain and the inlet to the converging portion of the exit nozzle. Test e rings with thrusts up to 960 N were conducted with gaseous oxygen and hydroxylterminated polybutadiene solid fuel. Average fuel regression rates up to seven times larger than those in similar classical hybrids were measured. Empirical correlations were developed to describe accurately the experimental regression rates over more than an order of magnitude variation in mass e ux. In addition to local mass e ux and oxygeninjectionvelocity,geometricenginevariables,suchasenginecontractionratioand length-to-diameterratio, had a signie cant ine uence on the measured regression rates. Nondimensional regression rate and heat transfer correlations were also developed. Throttling and restart capability were demonstrated. Nomenclature A = cross-sectional area, cm 2 B = blowing parameter cp = isobaric specie c heat, J/kg ¢K D = port diameter, cm Ea = activation energy, kcal/mole G = local mass e ux, kg/m 2 ¢s Ginj = injection mass velocity, kg/m 2 ¢s GO = gaseous oxygen mass e ux, kg/m 2 ¢s

The Bidirectional Vortex. Part 1: An Exact Inviscid Solution

In this paper, we derive an exact solution that describes the bulk fluid motion of a bidirectional coaxial vortex appropriate of a liquid propellant combustion chamber. The study is prompted by the need to characterize the flow inside a laboratory-scale thrust chamber. This chamber has the advantage of confining mixing and combustion to an inner vortex tube that remains separated from the chamber walls by an outer stream of swirling, low temperature oxidizer. Our mathematical model is based on steady, rotational, axisymmetric, incompressible, and inviscid flow conditions. In contrast to other studies of columnar vortices (where the axial dependence is ignored), our model accounts for the chamber’s finite body length. In fact, it incorporates the proper inlet and head-end flow conditions associated with a bipolar swirl-driven combustor. Based on the exact solution, several important flow attributes are illuminated. Among them is the location of the nontranslating vortex layer known as the mantle. This cylindrical layer separates the outer and inner vortex tubes (i.e., the updraft and the downdraft) and is confirmed using computational fluid dynamics and flow visualization.

The Bidirectional Vortex. Part 2: Viscous Core Corrections

This article focuses on the viscous core of the bidirectional flowfield arising in a swirldriven thrust chamber. By regularizing the momentum equation in the tangential direction, the boundary layer equation that controls the forced vortex near the chamber axis is obtained. After identifying the coordinate transformation needed to resolve the rapid changes near the core, an inner expansion is arrived at. This expansion is then matched with the outer solution associated with the free vortex; the latter is known to prevail in the outer region. By combining inner and outer expansions, uniformly valid approximations are obtained for the swirl velocity, vorticity, and pressure. These are shown to be strongly influenced by a dynamic similarity parameter that combines the mean flow Reynolds number and the chamber aspect ratio. Referred to as the vortex Reynolds number V, this dimensionless grouping enables us to quantify the characteristic features of the bidirectional vortex. Among them is the thickness of the viscous core which is found to decrease with the square root of V. The converse can be said of the maximum swirl velocity. In the same vein, the angular frequency of the rigid-body rotation of the forced vortex near the core is found to be linearly proportional to V. The form of the swirl velocity is reminiscent of the Burgers vortex; here, it is based on the aspect ratio of the thrust chamber. The resulting theoretical predictions compare favorably with experimental measurements and computational results over the length of the chamber.

Experimental investigation of a vortex-driven high-regression rate hybrid rocket engine

An experimental investigation to characterize the solid-fuel regression rate and combustion behavior of a vortex-driven, highregression rate, lab-scale hybrid rocket engine was performed. Gaseous oxygen was injected through a ring of tangential ports located between the end of the fuel grain and the exit nozzle. HTPB was the primary fuel of interest for regression rate studies. Transparent PMMA fuel grains were used for hot-firing visualization tests. The visualization tests showed that the tangential injection technique generated a pair of co-axial, co-spinning vortices in the cylindrical combustion chamber. Radial pressure measurements and the recovered fuel samples also indicated the presence of vortex flow. The swirling flow field seemed to augment the combustion process, and was effective in promoting high regression rates. Within the limited range of test conditions, regression rates 640% faster than those found in similar classical hybrids were demonstrated. Dimensional and non-dimensional correlations were developed to describe the average regression rates as a function of operating conditions. The regression rates showed a strong dependency on GOX mass flux, injection mass flux, port diameter, and grain length. The local regression rates were quite uniform for a particular distance from the injector, then began to decrease with axial location. Mixture ratio control was demonstrated by axially injecting a small amount of GOX through the motor head-end using an independent supply system.

Hot Flow Model of the Vortex Cold Wall Liquid Rocket

This study details the reactive flow simulation of the NASA sponsored cold wall bidirectional vortex chamber (CWBVC). The main features of the CWBVC are discussed in a former article in which the flow was treated as incompressible (see Fang, D., Majdalani, J., and Chiaverini, M. J., “Simulation of the Cold Wall Swirl Driven Combustion Chamber,” AIAA Paper 2003-5055, July 2003). The current reactive flow model is based on a two-stage, choked nozzle approach. Knowing that the nozzle is choked at the throat under normal operation, the CWBVC chamber is decomposed into two parts: the first part extends from the head end to the nozzle throat while the second extends from the throat to the nozzle exit plane. In the first simulation stage, the incompressible flow model is applied with the divergent part of the nozzle truncated out. In the second stage, compressibility is superimposed starting with the output from Stage I. This two-stage simulation reduces CPU time and helps to achieve convergence of the compressible, reactive flow model. Our simulations rely on a commercial CFD solver to handle the steady-state, three-dimensional, Navier-Stokes equations. Specifically, we employ a 3D, segregated, implicit, absolute scheme in conjunction with the realizable κ e − turbulence model of Launder and Spalding. Reaction mechanisms are simulated using the non-premixed combustion model with the adiabatic PDF look-up tables. The eight conventional chemical species are used in simulating hydrogen-oxygen combustion; these include: 2 O 2 2 2 2 2 and . At the conclusion of the reactive flow simulation, the existence of a bidirectional flow is demonstrated and the spatial invariance of the so-called mantle (that separates inner and outer vortex regions) is discussed. The cold wall effect is clearly extrapolated from the temperature maps over the combustion chamber wall. At the nozzle exit, shock waves are detected due to over-expansion. , H , H O, HO , H O , O, H OH

Development and testing of a vortex-driven, high-regression rate hybrid rocket engine

ORBITEC has been developing a new type of hybrid rocket under NASA MSFC Phase I and II SBIRs where the gaseous oxidizer is injected tangent to the inner fuel grain surface either through or at the bottom of the fuel grain wall. The resulting combustion chamber flow field is a bi-directional co-axial vortex where the spinning flow migrates up the outer wall to the head end of the engine, flows inwards towards the center of the engine, and out the engine nozzle. The motivations for developing this unique hybrid are that it regresses over 8 times faster than a conventional head end injected hybrid at a comparable oxidizer mass flux. Additionally, the flow field increases the effective combustion stay time for a given size chamber volume and nozzle throat area. It is expected that these two characteristics will allow increased hybrid performance, the use of a simple single port grain geometry, and increased grain case volumetric loading efficiency. It has also been discovered that the regression rate of this vortex hybrid can be tailored by modifying the injector geometry. Gaseous oxidizer would be used in flight versions of the engine where the full flow staged combustion cycle or expander cycle can gasify and deliver the oxygen. The experimental results of the Phase I effort are presented here. A total of 32 firings were conducted to investigate the effects of scaling, injector pattern, injector diameter, and solid fuel type. The vortex flow field was characterized through smoke testing, oil streak testing, and filtered video data of the firings. The fuel regression rate was found to be strongly dependant upon the injector geometry, oxidizer mass flux, and the corresponding strength of the vortex. Statistical analysis was used to develop a vortex regression law that predicts the regression rate as a function of these parameters. The Phase IISBIR is currently being performed where a 500 Ibf engine will be designed and tested.

Hybrid Rocket Investigations at Penn State University's High Pressure Combustion Laboratory: Overview and Recent Results

The Pennsylvania State University has actively pursued hybrid rocket fuel regression rate and combustion research for 15 years. Initial work focused on developing and testing a high-pressure slab-burning hybrid motor with X-ray diagnostics for characterizing the local, real-time regression rates of hydroxyl terminated p olybutadiene (HTPB) with gaseous oxygen. Fuel decomposition and pyrolysis investiga tions were also carried out. Follow-on work is continuing to investigate high regression r ate hybrid fuels with various metal additives in center-perforated hybrid motors using both HTPB and paraffin binders. The addition of aluminum powders to paraffin-based solid-fuel formulations was shown to increase the regression rates by a factor of 4 comp ared to neat HTPB. This regression rate increase corresponds to a mass-burning rate increase ~7 times that of HTPB when the increase in fuel density is considered. This artic le reviews some of the more significant results from previous investigations and presents r ecent data from on-going test programs. Nomenclature

Preliminary CFD analysis of the vortex hybrid rocket chamber and nozzle flow field

A laboratory scale hybrid rocket engine of a unique new design was found to achieve significantly higher fuel regression rates than in classic hybrids operating at similar conditions with similar propellants. Gaseous oxygen was injected through a ring of tangential ports located between the end of the fuel grain and the exit nozzle. A CFD analysis was employed to confirm the overall features of the experimentally observed flow field in the engine. A coaxial, co-swirling vortex flow was shown to develop in the cylindrical combustion chamber. The outer vortex spirals upward along the fuel surface, toward the forward end of the combustion chamber. The flow migrates radially inward as it approaches the head-end of the grain and develops into an inner, downward-spiraling vortex that exits the nozzle. This unusual flow field seemed to augment combustion, and was effective in promoting high regression rates. Regression rates eight times faster than classic hybrids were experimentally demonstrated using both PMM and HTPB fuel grains. A numerical investigation was conducted to determine the influence of design parameters, such as engine contraction ratio and injection velocity on flow-field characteristics. The swirling flow increases the gas velocity, and therefore the convective heat transfer, near the fuel surface. The serpentine flow path also allows for extended travel distance in the engine, which should increase combustion efficiency.

Fuel decomposition and boundary-layer combustion processes of hybrid rocket motors

Using a high-pressure, two-dimensional hybrid motor, an experimental investigation was conducted on fundamental processes involved in hybrid rocket combustion. HTPB (Hydroxyl-terminated Polybutadiene) fuel cross-linked with diisocyanate was burned with GOX under various operating conditions. Large-amplitude pressure oscillations were encountered in earlier test runs. After identifying the source of instability and decoupling the GOX feed-line system and combustion chamber, the pressure oscillations were drastically reduced from +/-20% of the localized mean pressure to an acceptable range of +/-1.5% Embedded fine-wire thermocouples indicated that the surface temperature of the burning fuel was around 1000 K depending upon axial locations and operating conditions. Also, except near the leading-edge region, the subsurface thermal wave profiles in the upstream locations are thicker than those in the downstream locations since the solid-fuel regression rate, in general, increases with distance along the fuel slab. The recovered solid fuel slabs in the laminar portion of the boundary layer exhibited smooth surfaces, indicating the existence of a liquid melt layer on the burning fuel surface in the upstream region. After the transition section, which displayed distinct transverse striations, the surface roughness pattern became quite random and very pronounced in the downstream turbulent boundary-layer region. Both real-time X-ray radiography and ultrasonic pulse-echo techniques were used to determine the instantaneous web thickness burned and instantaneous solid-fuel regression rates over certain portions of the fuel slabs. Globally averaged and axially dependent but time-averaged regression rates were also obtained and presented.