Rocket Performance Research Articles

レーザー着火マイクロ固体ロケットにおける燃焼室圧力制御 による性能向上

レーザー着火マイクロ固体ロケットにおける燃焼室圧力制御 による性能向上

Air-breathing performance of microwave rocket with reed valve system

The microwave rocket is a candidate for use as a future low cost launch system. Plasma generated by a millimeter-wave beam drives a detonation wave in a thruster tube, and the vehicle acquires impulsive thrust by exhausting the gas, similarly to a pulse detonation engine. The air breathing rate is important for thrust performance in multi-pulse propulsion. Therefore, an air-breathing mechanism using reed valves is under development. In this study, the air flow rate through the reed valves was computed and its performance was evaluated in terms of the partial filling rate: PFR. Results show that, when using reed valves, a 10–15-fold increase in thrust is obtainable over conventional performance.

신뢰성 보정된 발사체 개발비용 추정방안 연구

본 연구에서는 발사체의 개념설계 단계 동안 성능 변수뿐만 아니라 신뢰성에 따른 개발비용 추정방안을 분석하고자 하였다. 과거에는 발사체의 개발비용을 예측하기 위해서 주로 발사체의 성능과 건조 질량 등을 독립변수로 하는 추정 방법이 많이 이용되었다. 이러한 접근 방법은 비교적 근사하였지만 신뢰성 수준에 따른 비용 변화를 반영할 수 없기 때문에 원하는 신뢰성을 가진 발사체의 개발비용을 근접하게 예측하기는 쉽지 않았다. 본 논문에서는 이러한 비용추정 방법의 한계를 개선하기 위해 성능과 질량을 기반으로 한 비용 모델인 TRANSCOST에 신뢰성 개념을 도입하여 발사체 비용을 추정할 수 있는 방법론을 제시하였다. 이를 기반으로 신뢰성에 따른 한국형발사체(KSLV- II)의 개발비용을 추정했다. The purpose of this study is to perform the reliability-corrected development cost estimation of the launch vehicle at the conceptual design phase. In order to estimate the launch vehicle development cost, the estimation method based on the independent variable such as the rocket performance and dry mass has been mainly implemented up to now. This approach has made the approximate cost estimation possible, however, the cost variation according to the reliability requirement could not be reflected. In this paper, the cost estimation methodology that introduces the reliability factor in addition to the performance and mass in the TRANSCOST model is presented in order to improve the limitation of current cost estimation method. The development cost of KSLV(Korea Space Launch Vehicle)-II is estimated on the basis of this newly implemented concept with reliability as an added parameter.

Optimization of Hybrid Sounding Rockets for Hypersonic Testing

The use of ground-launched hybrid rockets to bring subscale models to flight conditions within the hypersonic flight corridor is investigated. The analysis is based on a multidisciplinary optimization procedure that couples the direct optimizations of the parameters that rule the engine geometry and operation and the indirect optimization of the trajectory. The final Mach number is maximized for given initial conditions and assigned payload and final altitude. Different rocket configurations are considered; in particular, the performance of single-stage and two-stage rockets are determined for different payload fractions. Results show that hybrid motors are suitable to accomplish this kind of mission, and that single-stage rockets are preferable. Only when the payload fraction is very small and large final velocity are sought, two-stage rockets may offer better performance.

Mathematical model and experimental results for cryogenic densification and sub-cooling using a submerged cooling source

Among the many factors that determine overall rocket performance, propellant density is important because it affects the size of the rocket. Thus, in order to decrease the size of a rocket, it may be desirable to increase the density of propellants. This study analyzes the concept of increasing the propellant density by employing a cooling source submerged in the liquid propellant. A simple, mathematical model was developed to predict the rate of densification and the propellant temperature profile. The mathematical model is generic and applicable to multiple propellants. The densification rate was determined experimentally by submerging a cooling source in liquid oxygen at constant positive pressure, and measuring the time rate of change in temperature with respect to vertical position. The results from the mathematical model provided a reasonable fit compared to experimental results.

Hybrid Rocket Performance Prediction with Coupling Method of CFD and Thermal Conduction Calculation

Hybrid Rocket Performance Prediction with Coupling Method of CFD and Thermal Conduction Calculation

Design and optimization of variable thrust hybrid rocket motors for sounding rockets

Besides safety and low-cost, the start/shutdown/restarting and throttling ability are the other two significant advantages of hybrid rocket motors (HRMs) compared with liquid and solid ones. In this study, a two-stage variable thrust and non-toxic 98%HP/HTPB hybrid rocket motor (VTHRM) is designed and applied in a sounding rocket, and the design parameters of the motor are analyzed and optimized. A computational program is developed to design the motor system structure, to predict the interior ballistics and the ballistic trajectory. A star grain and a wheel grain are compared. The design of experiment (DOE), variance analysis and the main effect analysis are employed to investigate the influence of the main design parameters on motor performance. The multidiscipline feasible (MDF) approach is applied to establish the optimization procedure after analyzing the system design structure matrix. A modified differential evolution algorithm is employed to maximize the load mass. The results indicate that the wheel grain could obtain a larger load mass and a lower length to diameter ratio, and that throttling markedly meliorates the motor and rocket performance. The conclusions drawn from the analysis and optimization could provide instructive guide and theoretical basis for engineering designs.

Development of preliminary design program for combustor of regenerative cooled liquid rocket engine

An integrated program was established to design a combustor for a liquid rocket engine and to analyze regenerative cooling results on a preliminary design level. Properties of burnt gas from a kerosene-LOx mixture in the combustor and rocket performance were calculated from CEA which is the code for the calculation of chemical equilibrium. The heat transfer of regenerative cooling was analyzed by using SUPERTRAPP code for coolant properties and by one-dimensional correlations of the heat transfer coefficient from the combustor liner to the coolant. Profiles of the combustors of F-1 and RS-27A engines were designed from similar input data and the present results were compared to actual data for validation. Finally, the combustors of 30 tonf class, 75 tonf class and 150 tonf class were designed from the required thrust, combustion chamber, exit pressure and mixture ratio of propellants. The wall temperature, heat flux and pressure drop were calculated for heat transfer analysis of regenerative cooling using the profiles.

Research on the Performance Detecting System of High-Precision Small-Scale Solid Rocket Motor

According to the performance detecting requirements in development and manufacture of solid rocket motor, a test platform had been set up based on flexible test fixture, a measurement & control platform had been built using virtual instrument technology, a set of solid rocket motor ground test performance detecting software including parameter calibration module, data measurement module and data processing module had been developed based on the LabVIEW7.1, and the solid rocket motor performance detecting system was constructed. The application results through testing normative solid rocket motor show that: the system could actualize parameter calibration expediently, experimental data collection and processing rapidly, even reach the measurement precision by 0.3% that required in project and could monitor the working process of solid rocket motor simultaneously, which satisfies high-precision requirement of motor performance detecting.

Preliminary analysis of hybrid rockets for launching nanosats into LEO

This work determines the preliminary mass distribution of hybrid rockets using 98% H2O2 and solid paraffin mixed with aluminum as propellants. An iterative process is used tocalculate the rocket performance characteristics and to determine the inert mass fractionfrom given initial conditions. It is considered a mission to place a 20 kg payload into a 300 km circular equatorial orbit by air launched and ground launched hybrid rockets usingthree stages. The results indicate total initial masses of about 7800 kg for a groundlaunched hybrid rocket and 4700 kg for an air launched hybrid rocket.

PENGARUH KETIDAKLURUSAN DAN KETIDAKSIMETRISAN PEMASANGAN SIRIP PADA PRESTASI TERBANG ROKET RX-250-LPN

The performance of a rocket depend on several factor, one of which is the value of drag experienced during flight. This drag is contributed by the shape and the many components of the rocket. This paper discussed the fin misalignment and asymetrisness in respect to its main axis, which results in a change in the drag of RX-250-LPN rocket. The discussion was conducted with assuming the variation of this fin alignment from 1 degree to 2 degree. DATCOM was used to obtain aerodynamic parameters of the rocket. In addition, MATLAB 6.3 was used to calculate the performance of rocket. The knowledge of this fin misalignment can be used to get the minimum accuracy needed for better fin setting during rocket assembly.

Conceptual design of wrap-around-fin rockets using the MDO methodology

The conceptual design of rockets is a system engineering that involves configuration, aerodynamics, engine, dynamic stability, and coupling effects. To obtain benefits from the synergistic effects, the latest research achievements, and the most powerful analysis tools efficiently, this paper focuses on how to apply the multidisciplinary design optimization (MDO) technology to the conceptual design of the long-range, large length to diameter ratio wrap-around-fin (WAF) rockets. The software iSIGHT was employed as the MDO framework, the multidisciplinary feasible (MDF) method as the MDO architecture, and the multi-island genetic algorithm combined with sequential quadratic programming as the search algorithm. By maximizing the payload ratio, the structural analysis, aerodynamics, engine, trajectory, and dynamic stability were considered comprehensively. A case study demonstrated that MDO could improve the rocket performance effectively.

Performance analysis of the pulse detonation rocket engine based on constant volume cycle model

As the propagation time of detonation is very short in the actual operation process of a pulse detonation rocket engine (PDRE), the gases expelled slightly from the detonation chamber in this process is negligible. Therefore, a mathematical model of constant volume cycle (CVC) can be used to approximately depict the operation process of PDRE. Based on CVC model, a systematic analysis of PDRE’s performance is conducted in this paper. Combined CVC model with Chemical Equilibrium and Applications (CEA) code, the specific impulse of PDRE using different propellants could be obtained. Utilizing this model, the specific impulse of kerosene/oxygen PDRE in vacuum and at finite backpressures is estimated. When the backpressure exists, there is an optimum nozzle area expansion ratio in the nozzle design of PDRE, similar to conventional steady-state rocket. The performance of PDRE and conventional steady rocket is compared respectively, under the two conditions of the same supply system and the same maximum material stress. When PDRE and conventional steady rocket have a comparable performance, the filling pressure of PDRE is much lower than that of steady rocket.

Investigation of Rocket Exhaust Diffusers for Altitude Simulation

The design and operational parameters of rocket exhaust diffusers equipped to simulate high-altitude rocket performance on the ground were investigated and characterized using a comprehensive approach (theoretical, numerical, and experimental). The physical model of concern includes a rocket motor, a vacuum chamber, and a diffuser, which have axisymmetric configurations. Further, the operational characteristics of a rocket exhaust diffuserwereanalyzed froma flowdevelopmentpointof view.Emphasiswasplacedondetailed flowstructure inthe diffuser, to observe the pressure oscillation in both the vacuum chamber and diffuser, which determines the minimum rocket-motor pressure required to start the diffuser. Numerical simulations were compared with experimental data on startup and in operational conditions to understand the effects of major design parameters, including the area ratio of diffuser to rocket-motor nozzle throat, the rocket-motor pressure, and the vacuumchamber size. Nomenclature Ad = inner cross-sectional area of diffuser Ade = exit cross-sectional area of diffuser Ae = exit cross-sectional area of rocket nozzle At = throat cross-sectional area of rocket nozzle

BORON NANOPARTICLE-RICH FUELS FOR GAS GENERATORS AND PROPELLANTS

This study reports on the production and characterization of energetic polymer-capped boron for solid fuel applications. It is known that the addition of metal and metal-like particles to solid fuels and propellants can improve the performance of both rocket and air-breathing propulsion systems. The use of boron is very attractive for these applications due to its high heat of combustion on both a gravimetric (58 kJ/g) and volumetric (136 kJ/cm3) basis. However, the exploitation of the high theoretical energy of boron has been limited by a few undesirable properties of this metal. Among them, one notes the existence of a resilient oxide layer on the particle surface affecting the ignition and combustion of boron particles. The capping of boron nanoparticles with a polymer can provide a solution to the aforementioned problem. The use of an energetic polymer for that purpose can bring additional heat close to the surface of the nanoparticles and facilitate their ignition. In the present work, boron nanoparticles were produced by synthesis of surface-functionalized boron. The route was adapted to obtain additional hydroxyl-functional groups on the particles. These hydroxyl sites were used to graft a diisocyanate, and then produce an energetic polymer matrix based on polyurethane chemistry by addition of glycidyl azide polymer, resulting in boron nanoparticles coated by energetic polymers. This can lead to significantly enhanced boron particle combustion.

Hydrodynamics and Thrust Characteristics of a Water-Propelled Rocket

An undergraduate experiment has been developed to measure the performance of a pressurized water rocket and compare test data with an analytical model developed from fluid momentum. A test stand allows the net thrust of the rocket to be measured, as well the air pressure and temperature inside the rocket, as functions of time. The model compares well with test data from four conditions (combinations of two initial water heights and two initial air pressures). The air in the rocket is assumed to undergo polytropic expansion, and a value of 1.1 for the polytropic exponent is found to best fit the model to experimental conditions. Experimental observations reveal that at higher air pressures the air mixes with the expelled water, resulting in a two-phase flow that reduces the net thrust of the rocket. A pedagogical approach to the experiment is described in detail.

Thrust Performance of Microwave Rocket Under Repetitive-Pulse Operation

An experiment was conducted on a microwave rocket (repetitive-pulse millimeter-wave-beam-powered thruster) with repetitive pulses. A thruster model with a forced-breathing system was used. The forced-breathing system supplies fresh air from the thrust wall into the thruster. The pressure histories in the thruster weremeasured and the propagation velocity of the shock wave and the thrust impulse were deduced. Results show that although the propagation velocity was identical to the result for the single-pulse operation at the first pulse, the propagation velocity increased after the second pulse. Similarly, the impulse decreased after the second pulse. The dependence of the propagation velocity of the shock wave and the thrust performance on the partial-filling rate of the fresh air was comparedwith that of the thrust-generationmodel with a forced-breathing system. The experimental results showed good agreement with those obtained using the model.

Flow dynamics at the minimum starting condition of a supersonic diffuser to simulate a rocket’s high altitude performance on the ground

A numerical analysis was conducted to investigate and characterize the unsteadiness of the flow structure and oscillatory vacuum pressure inside of a supersonic diffuser equipped to simulate high-altitude rocket performance on the ground. A physical model including a rocket motor, vacuum chamber, and diffuser, which have axisymmetric configurations was employed. Emphasis was placed on investigating the physical phenomena of very complex and oscillatory flow evolutions in the diffuser operating very close to the starting condition, i.e. at a minimum starting condition, which is one of the major important parameters from a diffuser design point of view.

Thrust Termination Dynamics of Solid Propellant Rocket Motors

A CTIVE thrust termination or reversing of a separatingmotor are ways to facilitate the staging process and improve the performance of a multistage rocket [1–4]. This is done by opening a secondary nozzle located on the motor head and causing rapid depressurization of the motor chamber. During a rapid change in the chamber pressure the burning rate becomes a nonlinear function of pressure and its time rate of change. The strong dependence of the solid propellant burning rate on the chamber pressure results in a transient nonlinear behavior that may cause premature dynamic extinguishment of the motor. Solid propellants’ transient burning and extinguishment during rapid pressure change have been the subject of a number of studies and several transient burning models have been developed [5–8]. Zeldovich has proposed a comprehensive model based on the unsteady rate of heat transfer to the propellant surface and the time rate of change of the temperature distribution in the solid propellant [6]. This model requires specification of many experimentally determined parameters. The transient burning rate model proposed by Von Elbe and McHale is easy to implement and requires instantaneous pressure and its time rate of change for the regression rate prediction [5]. References [1,2] studying motor transient behavior have used a control volume approach for the gas dynamic simulation of the motor’s internal ballistics. However during the active thrust termination (or reversing) process there is a considerable pressure variation along the motor port and there is a rapid propagation of expansionwaves, through the thrust termination nozzles. These expansion waves play an important role in the transient thrust behavior and a more comprehensive flow model should be used for the internal ballistics’ simulation.

Influence of a Conical Axial Injector on Hybrid Rocket Performance

This paper was aimed at analyzing the static engine firings results obtained by means of a hybrid rocket where gaseous oxygen was supplied into axial-symmetric polyethylene cylindrical grains through two different injector configurations: an axial conical subsonic nozzle and a radial injector. The axial injector is considered interesting because of its easy design and the remarkable feature that it produces no significant pressure oscillations for the stabilizing effect due to the hot gas recirculation zone established within the combustion port. To take advantage of its qualities, the assessment of the regression rate variations under the flow field generated by this configuration is required. For the investigated set of operating conditions, the instantaneous regression rates exhibit a time-and-space dependence caused by the impinging jet zone dynamics, while the average regression rates are higher and less mass flux dependent than those achieved with the radial injection motor and expected from the turbulent flow through pipes. A comparison to the data from the radial injector was further drawn in terms of combustion efficiency, fuel consumption profiles, and combustion stability. The radial injector, at the same mass flux and pressure, produces lower regression rates, high pressure oscillations and worse combustion efficiency at the same L*, but more uniform fuel consumption.