Propeller Parameters Research Articles

Four Quadrants Numerical Prediction for Hydrodynamic Performance of Open-Water Propeller

In order to obtain the four quadrants hydrodynamic performance of open water propeller by means of CFD, a mathematic model of three dimensional coordinates points was established and programmed using Matlab based on the basic parameters of propeller. A smooth model propeller was made by importing these points into front end software. Then taking AU model for example, numerical simulations of propeller turning ahead while going ahead, turning ahead while going astern, turning astern while going ahead and turning astern while going astern were carried out. At the same time, the thrust and torque coefficients were presented. The simulation results showed good agreement with the results of tank test. The influence of mesh generation and computational domain on open-water performance were also discussed.

Experimental Studies of Propellant Loading Parameters and Plasma Flow-Field Interactions

Propellant-plasma interaction has been attracting great interests among most researchers for application to an electrothermal-chemical (ETC) gun. Previous works have shown that propellant-plasma interaction can enhance burning rates due to increased surface erosion of propellants as compared to that of conventional ignition. Plasma radiation and heat convection are the dominant mechanism of energy transfer based on both experimental data and numerical simulation. However, little literatures were reported about a detailed understanding of propellant loading parameters and the plasma flow-field distribution interactions, which is very important in deciding a successful propellant-charge design in ETC guns. The pressure-time curves during interrupted-burning process of different propellants such as triethyeneglycol dinitrate and single-base propellants ignited by plasma were recorded by pressure sensor in interrupted-burning setup. Influence of loading parameters, including loading density, distance between propellants and plasma jet, and propellant holding methods at the same selected location in plasma flow-field on plasma-ignition performance of propellants was discussed based on experimental results. As the loading density of propellants increases, the porosity of the propellant bed becomes small. Moreover, the penetrating depth of plasma into propellant bed shortens. Therefore, the plasma energy attenuates rapidly, and the ignition-delay time of propellants becomes longer. The ignition-delay time is longer when the distance between propellants and plasma jet increases. At the same selected location in plasma flow-field, different holding methods of propellant burning surface to plasma jet have less influence on the ignition, combustion process of propellants, and the ignition-delay time. The matching test results on propellant loading parameters and plasma flow-field distribution interaction play a major role in improving charge configurations at chosen plasma flow-field and understanding plasma-propellant interaction better.

A Hybrid Optimization Approach for SRM FINOCYL Grain Design

This article presents a method to design and optimize 3D FINOCYL grain (FCG) configuration for solid rocket motors (SRMs). The design process of FCG configuration involves mathematical modeling of the geometry and parametric evaluation of various independent geometric variables that define the complex configuration. Virtually infinite combinations of these variables will satisfy the requirements of mass of propellant, thrust, and burning time in addition to satisfying basic needs for volumetric loading fraction and web fraction. In order to ensure the acquisition of the best possible design to be acquired, a sound approach of design and optimization is essentially demanded. To meet this need, a method is introduced to acquire the finest possible performance. A series of computations are carried out to formulate the grain geometry in terms of various combinations of key shapes inclusive of ellipsoid, cone, cylinder, sphere, torus, and inclined plane. A hybrid optimization (HO) technique is established by associating genetic algorithm (GA) for global solution convergence with sequential quadratic programming (SQP) for further local convergence of the solution, thus achieving the final optimal design. A comparison of the optimal design results derived from SQP, GA, and HO algorithms is presented. By using HO technique, the parameter of propellant mass is optimized to the minimum value with the required level of thrust staying within the constrained burning time, nozzle and propellant parameters, and a fixed length and outer diameter of grain.

Design calculation of superlight piston-engined aircraft flight characteristics

Algorithms for determining fixed-pitch propeller parameters and some flight characteristics of a superlight aircraft at the early stage of designing are described.

Propeller Geometric Parameter Extraction From Inspection Data Clouds

Methods and conventions for propeller geometric parameter extraction from inspection data clouds are presented. Steps in the extraction process include constructing a representation of the inspected propeller from inspection data clouds, constructing sections through the representation, mapping the sections from three to two dimensions, and computing propeller parameters for the sections.

Temperature and Pressure Sensitivities of Burning Wave Parameters of Nitramine-Containing Propellants and HMX

Burning rate and burning surface temperatures have been obtained experimentally for modern double-base propellants containing nitramine additives and for HMX. Temperature sensitivities of burning wave parameters were estimated from the experimental data measured at pressures of 20 and 100 atm and at sample temperatures of i 80, +20, +100 C. Evaluation of standard deviations of the estimations shows that smoothing procedures for dependencies of burning wave parameters on pressure and temperature are necessary to obtain correct values of thesensitivities. Calculations ofthecriteria forstable combustion show that Novozhilov’ s [Novozhilov, B. V., “ Nonstationary Combustion of Solid Rocket Fuels,” Nauka, Moscow, 1973 (Translation AFSC FTD-MD-24-317-74) (in Russian) and Novozhilov, B. V., “ Theory of Nonsteady Burning and Combustion Stability of Solid Propellants by Zel’ dovich‐ Novozhilov Method,” Nonsteady Burning and Combustion Stability of Solid Propellants , edited by L. DeLuca, E.W. Price,and M. Summere eld, Vol. 143,Progressin Astronauticsand Aeronautics,AIAA,Washington, DC, 1992, Chap. 15, pp. 601‐ 641] criterion correctly predicts regimes of stable combustion of the propellants and HMX. An analysis of burned surface irregularities of the propellants cone rms one-dimensional character of the combustion under the investigated conditions.

Reply to comments by A.R. Shouman on “kinetic parameters of composite propellants from thermogravimetric data” by V.K. Rao et al

We present an experimental study of the “tulip flame” phenomenon using high-speed photography. Contrary to most previous studies, the work is in a simplified quasi-constant pressure configuration in a half-open tube. It is shown that the salient features of the different stages of the flame propagation and shape can be explained by a simple geometrical model of the interaction between the flame front and the gas dynamics. In particular, the tulip flame results from an inversion of the flame front curvature caused by the deceleration related to loss of flame surface area. Finally, the experimental results obtained by other authors in closed vessels are in reasonable agreement with the analysis presented.

The Bootstrap Approach to Predicting Airplane Flight Performance

The Bootstrap Approach (TBA) to fight performance is, for fixed-pitch propeller-driven airplanes, a set of simple formulas in nine aircraft parameters-four from brief flight tests-plus the variables of aircraft weight and air density. The formulas are derivable from the empirically supported linearity of the propeller polar. TBA outputs include five full-throttle or gliding V-speeds (Vx Vy Vm Vbg Vmd) and a common version of long range cruise speed (VLRC) as well as, for any air speed, thrust, drag, rate of climb (thereby service ceiling) or descent, and flight path angle. Knowledge of engine speed (RPM) and fuel consumption rate allows calculation of (partial-throttle) speeds Vbr for best range and Vbe for best endurance. An extension of TBA to airplanes with constant-speed propellers substitutes two different propeller parameters and suggests a general propeller chart. Though based on empirical graphs rather than formulas, this extension easily incorporates partial-throttle operation.

Flight test of a half-scale unmanned air vehicle

A half-scale unmanned air vehicle was flight tested to provide lift and drag data for correlation with the fullscale vehicle. Additional work was carried out to determine if wing drag could be reduced with an improved surface finish and a trailing-edge modification. Ground tests for power and thrust using a torque stand and a lowspeed wind tunnel supported the flight tests for the determination of engine and propeller parameters. A panel method was used to predict the induced drag behavior of the tested air vehicle. Parasite drag was predicted by build-up methods. The predicted parasite drag for the half-scale fell short of flight-tested values by about 25%, but that predicted for the full-scale vehicle correlated well with the flight data. Within the scatter of data, the induced drag correlated satisfactorily. A comparison was made to lift and drag data extracted from full-scale, idlepower glide tests, and correlation was poor. Implications are that residual thrust prevents these glide-test data from accurately representing the full-scale drag polar.

Assessment of existing and future launch vehicle liquid engine development

Existing liquid propellant engines for large launch vehicles are described in terms of pertinent engine and propellant parameters and their launch vehicle application. The development approach and the maturity of engine technology which prevailed prior to and early in specific engine development programs are discussed including lessons learned. New engines, including improved conventional and new concepts that could support the next generation launch vehicle, are delineated with emphasis on technology. The technology maturity and development needed to alleviate the potential development risk are presented along with projected gains in performance, operability, reusability, reliability and producibility. A technology ranking methodology which incorporates a relative transportation system life cycle cost (LCC) as the payoff function is developed. The methodology is useful for establishing preliminary but timely cost effectiveness rankings of various technologies. The methodology uses conventional cost estimating relations (CER) in non-dimensional form. The relative overall transportation system payoff resulting from the implementation of new propulsion system technology is developed from recurring and non-recurring costs of major transportation system elements including the vehicle, operations and launch complex. A ranking of engine concepts and associated technologies is given for several transportation system candidates which serve a high activity mission model.

Energetics of Propellants, Fuels and Explosives; a chemical valence approach

AbstractA simple approach involving the chemical valences of the fuel and the oxidizer elements present in a combustible mixture is described to evaluate the energetics related parameters of propellants, fuels and explosives. It simplifies the stoichiometric balancing of complex combustion equations, and provides an easy method of calculating the elemental stoichiometric coefficient. The method correctly predicts whether a mixture is fuel‐lean, fuel‐rich or stoichiometrically balanced. The calorimetric value of various stoichiometrically balanced combustible systems has been shown to be linearly dependent upon their total oxidizing (or reducing) valences. This relationship has been used successfully to evaluate the calorific value of fossil fuels. For fuel‐rich explosives, a new valence dependent parameter has been derived which is found to be related with properties such as detonation velocity, heat of explosion and impact sensitivity.

Form Function for Propellants in Closed Vessel Work

AbstractA general theoretical cubic form function for use in closed vessel work during the evaluation of absolute ballistic parameters of propellants is introduced along with the expressions for the coefficients for a number of common propellant shapes. The conditions under which this form function is accurately applicable are also indicated. This form function is superior and gives much more accurate results than the quadratic form function which is in wide use at present. The differences of the cubic form function from the quadratic are also explained.

Evaluation of solid propellant ballistic properties

{fx193-02} have been used for deteminig the ballistic properties and {fx193-03}. The fires and older techniques consists of burning a small {fx193-04} under steady-state conditions. A somewhat newer technique consists od obtaining a {fx193-05} time traces by burning a small end burning propellant grain in a {fx193-06} under uncertaunty conditions. In the format technique, the data are correlated using the {fx193-07} a number of experiments at different steady-state pressures being recognised. For the sites techniques data from only one experiment are needed. The correlation is there made by comparing the first and second time derivatives of the pressure obtained from the pressure-{fx193-08} to as equations derived from combustion vessel parameters and the equation, {fx193-09}. While the data for mass proportions compare quite favonbly to the equation derived using the steady state burning equations, a better correlation is obtained by using an unsteady state equaions for the burning state. In addition, the use of the unsteady-state equation permits the determinations of a low-pressure deflegration limit below which the propellant would be difficult to ignite. The ow-pressure deflegration limit is found to to a function of the propellant parameters primarily including the theoretical differently of the enbursed propellant. For propellants having a small value of the theoretical differently, the values for the low-pressure deflagrtion limit would be correspondingly small.

A power prediction method and its application to small ships

A mathematical model is derived by which it is possible to describe the resistance and propulsion properties of ships as a function of the Froude number, the Reynolds number and the general form and propeller parameters. The coefficients of the mathematical model have been calculated for small ships such as trawlers and tugs from random tank data, obtained from the N.S.M.B.-statistics. With the present method it is possible to make an approximation of the resistance and required power which is accurate enough for design purposes. The method is suitable for programming.

A review of contemporary solid rocket motor performance prediction techniques

Solid rocket motor performance prediction including motor pressure, thrust and propellant parameters, discussing mathematical modeling of flow fields

Aluminized composite solid-propellant burning rates in acceleration fields

The average burning rates of a parametric series of propellants were determined in acceleration fields up to 1000 g using a combustion bomb mounted on a centrifuge. The propellant strands were burned at 500, 1000, and 1500 psia with the acceleration field directed normal and into the burning surface. Acceleration caused a time-dependent increase of as much as 100% in the burning rate of relatively slow-burning aluminized propellants. The relative amount of burning-rate increase for the various propellants was strongly dependent on the amount of aluminum retained on the propellant surface during burning. The burning rates of two very fast-burning propellants were found to be essentially accelerationindependent. OLID-PROPELLANT rocket motors are currently being used in applications which subject the metallized propellant grains to acceleration fields. The propellant burning rates increase appreciably when the acceleration field is imposed normal and into the burning surface.1'2 Postfire inspections of motors containing metallized propellants have revealed the presence of metal and/or metal oxide residue in the motor cases. The propellant burning-rate increase and the retention of residue appear to be interrelated. The review and abstracting of the literature on the effects of acceleration on solid-propell ant performance are presented in Refs. 3 and 4. Although progress has been made toward an understanding of some aspects of burning-rate augmentation in acceleration fields, it was believed that additional experimental studies with well-controlled propellant parameters would yield valuable insight. A parametric series of aluminized propellants was formulated to establish the importance of some of the propellant parameters that could be controlled to minimize the effects of acceleration. The objectives of this investigation were twofold. The first objective was to obtain the quantitative effect of acceleration on the burning rate of the propellants as a function of propellant composition, acceleration level, and time. The second objective was to obtain quantitative information on the amount of aluminum and/or aluminum oxide residue retained in the inhibitor case as a function of propellant -composition, acceleration level, pressure level, and propellant burning rate.

Monitoring the combustion process in large engines.

References 1 Valentine, R. S., Dean, L. E., and Pieper, J. L., An improved method for rocket engine performance prediction, J. Spacecraft Rockets 3, 1409-1414 (1966). 2 Crisman, P. A., Goldwasser, S. R., and Petrozzi, P. J., A general computer program for calculation of rocket performance parameters of propellants containing the atomic species, C, H, N, F and Cl, Proceedings of the ARS Propellant Thermodynamics and Handling Conference (American Rocket Society, New York, 1960), Special Rept. 12, pp. 293-313. 3 DeLeuze, N. C., Walsh, T. J., and Petrozzi, P. J., Calculation of chemical kinetic losses in a rocket nozzle, AerojetGeneral Corp., Thrust Chamber Engineering Rept. 9642:0046 (January 6, 1966). 4 Pieper, J. L., Dean, L. E., and Valentine, R. S., Mixture ratio distribution—its impact on rocket engine performance, Aerojet-General Corp., Thrust Chamber Engineering Rept. 9642: 0062/Revision B (December 21, 1966). 5 Rupe, J. H., A correlation between the dynamic properties of a pair of impinging streams and the uniformity of mixtureratio distribution in the resulting spray, Jet Propulsion Lab. Progress Rept. 20-209 (March 28, 1956). 6 Beltran, M. R., Breen, B. P., Kosvic, T. C., Sanders, C. F., Hoffman, R. J., and Wright, R. O., Liquid rocket engine combustion instability studies, Dynamic Science Final Report, Contract AF 04(611)-10542, AFRPL-TR-66-125 (July 1, 1966).

Transient effects in solid propellant burning

From theoretical considerations, stability criteria for solid propellant burning behavior are derived. The analysis considers the transient temperature field of the solid subjected to a cosine pressure perturbation in the gas stream. The results indicate that the modes of propellant burning behavior are influenced by the variation of the temperature field within the solid propellant. This effect is represented by the term C = [vto/ot]. The theory demonstrates that for different values of C, an initially stable propellant will burn in an unstable manner. In addition, the theory indicates a range in propellant parameters for which unstable burning always occurs.