Rocket Design Research Articles

Heat transfer measurement near injection hole of supersonic nozzle with a sonic jet injection

Heat transfer measurements were conducted to investigate surface heat transfer characteristics resulting from the interaction between a sonic jet and supersonic nozzle crossflow. A sonic jet was injected into the nozzle where the average Mach number is 2.88. Experiments were carried out for three different momentum flux ratios (J=0.5, 1.0, 1.5). A half-nozzle was introduced to measure heat transfer on the inner walls of the nozzle using Infrared thermography. Pressure measurements using PSP and oil flow visualization were also conducted to understand the heat transfer. From the oil flow visualization and heat transfer results, it was found that the area around the hole influenced by recirculation vortexes exhibited high heat transfer coefficients, showing up to 5 times higher values compared to the freestream. Furthermore, as the momentum flux of the jet increased, both the heat transfer augmentation area and the heat transfer coefficient also increased. To analyze the heat transfer characteristics between the nozzle and flat surface, comparisons were made using dimensionless heat transfer coefficients. The nozzle exhibited higher heat transfer in a smaller interaction area compared to that of a flat surface, but the average heat transfer coefficient near the hole was lower. Accordingly, distinct thermal treatment strategies should be employed for nozzles and flat surfaces, especially near injection holes. These insights could be valuable for the thermal design of rockets and other supersonic/hypersonic mobility systems.

Exploration of Students' Engineering Thinking Skills Through STEM Integrated PjBL Model Design Water Rocket Momentum Material

The research aims to analyse the engineering thinking skills of students by integrating the PjBL model with STEM water rocket design on momentum materials. This research is an experimental study. The design of this study was qualitative research. There are three stages to the research stage: preparation, implementation, and the final stage. The researchers collected data through observation, interviews, posttests, and documentaries. The focus of this study is on students' engineering thinking skills. The researchers interactively analyse the collected data during the phases of data collection, reduction, presentation, and conclusion. Additionally, the researchers conducted data credibility tests using triangulation techniques to assess the appropriateness of the research data. The research results predominantly classify the six indicators of engineering thinking as very good. Based on the findings and deductions, it is concluded that most students' engineering thinking abilities, when combined with the PjBL model and STEM water rocket design material, score highly on the six indicators of creative problem-solving. Students in this study experienced new knowledge and learning experiences that enabled them to solve daily problems by applying engineering thinking skills and physics concepts.

Home-Based Rocket Design and Fuel Project

This detailed research paper delves into the practicality and step-by-step process of creating a small-scale rocket and manufacturing its fuel from scratch within a home setting. The main objective of this study is to furnish a thorough manual for individuals passionate about amateur rocketry, emphasizing the importance of safety measures, materials, design concepts, and fuel composition. The project showcases the educational and experimental possibilities of rocketry in a domestic context, which can enhance knowledge of propulsion science and engineering fundamentals.

General Burnback Analysis for Anisotropic and Heterogeneous Solid Propellants

For several decades, researchers have successfully computed surface burnback to perform complex analyses and simulations related to the combustion of solid propellants. Many of these methods are heuristic or empirical, lacking a rigorous foundation, whereas others introduce numerical issues not related with the involved physics. This paper establishes a common theoretical and numerical core that bases the general study of burnback problems even in the case of designing nonconventional propellants. It starts with Piobert’s statement as a first principle, derives that the eikonal formulation holds for propellants with heterogeneous and anisotropic recession rates, and gives a general mathematical structure for both situations. Then, several configurations found by rocket designers are numerically solved in order to show the efficiency and the accuracy of the theory. The main conclusion is that a direct numerical integration of the eikonal equation using a time marching method is enough in terms of computational efficiency and accuracy to track the combustion surface of any anisotropic and heterogeneous solid propellant.

Numerical Study of Supersonic Jet Characteristics of Underwater Solid Rocket Engine

The structure of the rocket supersonic jet changes periodically when the solid rocket engines are ignited underwater. To investigate the flowfield structure and thrust characteristics of underwater gas jets emanating from solid rocket engines, this study establishes a numerical model for the multiphase flow of underwater jets based on the volume of fluid and realizable k-ε models. The effectiveness of the model is validated through comparison with experimental data. Subsequently, numerical simulations of tail jets from solid rocket engines are conducted at different water depths. The results indicate that during the underwater operation of solid rocket engines, the tail jet undergoes periodic processes such as expansion, constriction, rupture, and recoil. At larger depths, the environmental pressure exerts a more significant effect on the gas jet, thus resulting in a slower axial development of gas bubbles but more significant periodic evolution processes. Unlike the case in air, the thrust of solid rocket engines oscillates significantly when ignited underwater. In deeper water environments, the engine thrust presents a smaller oscillation amplitude. The findings of this study provide a theoretical basis for the design of solid rocket engines ignited underwater.

Small Solid-Model Rocket Design and Soft Landing Trajectory Planning

A recoverable rocket is one of the keys to reusable launch technology. Currently, most recoverable rockets are powered by liquid engines, which are costly and complex compared to solid engines. In this paper, we perform soft landing trajectory planning for the designed solid rocket. First, a small model rocket equipped with a landing engine is innovatively designed. Second, a six-degree-of-freedom motion model of the rocket is established, and a minimum-landing-error problem is formed based on the landing engine. Third, successive convexification algorithm is applied to solve it. In the algorithm, a method of quaternion constraint correction is proposed to avoid the infeasibility of the problem. Lastly, the correctness and robustness of the algorithm are verified via three simulations. The possibility of rapid trajectory generation is verified by code optimization on a high-performance digital signal processor.

Statement of requirements on the accuracy of rocket CFD analysis using exterior ballistics for example rocket models

In this paper, the problem of acceptable accuracy of a rocket’s aerodynamic characteristics obtained from CFD (computational fluid dynamics) analysis for defining its aerodynamic characteristics is considered from the point of view of exterior ballistics, particularly by the estimation of the rocket’s simulated flight range change caused by CFD errors. The results obtained in this paper allow stating the requirement on acceptable accuracy of defining the aerodynamic characteristics for developing new rockets through CFD analysis based on several examples, which can be propagated to other cases of rocket design and development. Using this study as a method, the CFD engineers who work on rocket design and development can estimate the accepted accuracy margin of their CFD analysis and research solutions to get reliable results for the expected flight characteristics of the rocket.

Aero-spike and aero-disk effects of on wave drag reduction of supersonic flow past over blunt body

PurposeThe purpose of this study is to explore the effects of aerospike and hemispherical aerodisks on flow characteristics and drag reduction in supersonic flow over a blunt body. Specifically, the study aims to analyze the impact of varying the length of the cylindrical rod in the aerospike (ranging from 0.5 to 2.0 times the diameter of the blunt body) and the diameter of the hemispherical disk (ranging from 0.25 to 0.75 times the blunt body diameter). CFD simulations were conducted at a supersonic Mach number of 2 and a Reynolds number of 2.79 × 106.Design/methodology/approachICEM CFD and ANSYS CFX solver were used to generate the three-dimensional flow along with its structures. The flow structure and drag coefficient were computed using Reynolds-averaged Navier–Stokes equation model. The drag reduction mechanism was also explained using the idea of dividing streamline and density contour. The performance of the aero spike length and the effect of aero disk size on the drag are investigated.FindingsThe separating shock is located in front of the blunt body, forming an effective conical shape that reduces the pressure drag acting on the blunt body. It was observed that extending the length of the spike beyond a specific critical point did not impact the flow field characteristics and had no further influence on the enhanced performance. The optimal combination of disk and spike length was determined, resulting in a substantial reduction in drag through the introduction of the aerospike and disk.Research limitations/implicationsTo predict the accurate results of drag and to reduce the simulation time, a hexa grid with finer mesh structure was adopted in the simulation.Practical implicationsThe blunt nose structures are primarily employed in the design of rockets, missiles, and re-entry capsules to withstand higher aerodynamic loads and aerodynamic heating.Originality/valueFor the optimized size of the aero spike, aero disk is also optimized to use the benefits of both.

Coherent solutions to roll damping derivatives evaluation for a generic rocket model

This paper presents a coherent approach to evaluate the roll damping derivatives for the standard Basic Finner Model. The study compares and analyses the results obtained through a range of techniques, including experimental testing, numerical simulations and semi-empirical models. The study aims to evaluate the reliability and accuracy of these methods and to identify the factors that contribute to their sensitivity. The paper concludes by summarizing the findings of the study and discussing the implications of the results for the design and operation of rockets. The experimental and numerical analysis used in this study provides a robust and comprehensive evaluation of the roll damping coefficient.

Numerical study on the effect of mass flux on the low frequency pressure oscillation occurred in oxygen jet condensation

AbstractIn the delivery pipeline of cryogenic liquid rockets, low‐frequency pressure oscillation always occurs with the oxygen jet condensation. It is a potential hazard to the normal operation of the propulsion system. It is necessary to study the characteristic of low‐frequency pressure oscillation precisely, especially the influence factors of frequency and amplitude. Aiming at the effect of mass flux on low‐frequency pressure oscillation, an oxygen injection condensation simulation is carried out based on a modified mass transfer model. The height function method is adopted to calculate the interfacial curvature, which is the key factor to capture the characteristic of pressure oscillation. The numerical result indicates that there are two typical oxygen jet condensation patterns under different mass fluxes: suck‐back chugging and slight pulsation. The amplitude of suck‐back chugging is 30–120 kPa, while that of slight pulsation is 1–3 kPa. When it comes to suck‐back chugging, the liquid oxygen suck‐back flow phenomenon occurs. The upper and lower boundaries are found to distinguish the suck‐back chugging and slight pulsation patterns, which are caused by the turbulent effect of liquid oxygen and the subcooling of liquid oxygen, respectively. The highest amplitude of pressure oscillation is about 114 kPa, appearing at the intersection of the upper and lower boundary. It is the most dangerous operation condition, which should be avoided in the design of the propulsion system. These conclusions provide sufficient theoretical guidance for the design of the cryogenic liquid rocket.

Comparative study of entropy distribution for generalized fluid between an inclined channel in the perspective of classical and non-Fourier’s law

The key challenge in the design of thermal appliances is to degrade entropy production with maximal energy dissipation to improve system efficiency in a variety of engineering applications, such as nozzle design,converging die, and rocket design. Due to the existence of both tangential and radial components, which have a significant impact on the fluid movement, coupled transport rates, and entropy development in an inclined channel presents a greater challenge. The current inquiry intended to address the combined impacts of Cattaneo-Christove (C-C) heat flux and classical theory on thermal dissipation, along with irreversibility distribution in a tilted porous media. The Carreau nanofluid confined between two finite intersecting plates in the presence of a porous Darcy-Forchheimer medium is examined. The C-C heat flux, variable thermal conductivity, and mass diffusivity are all used to investigate heat transport. The flow originates from a source located at the intersection of the plates and constitutes the momentum conservation. The investigation includes entropy and Bejan analysis. The premises guide the mathematical modelling of the flow field. Through bvp4c techniques, numerical solutions are attained. Converging section velocity increases in response to increased porosity and the Forchheimer number. Variation in temperature is significant when classical heat flux is reverted. Higher thermophoretic and porosity parameters result in a higher rate of entropy formation, while the Darcy parameter promotes drastic increase. Bejan is an escalatingfunction ofBrownian and Darcy parameters, while the porosity parameter causes it to fall. Porosity and the Forchheimer number regulate skin friction.

Space rocket for air-rocket system

The traditional and most used method of launching payloads into Earth orbit is to launch a carrier rocket from the surface. An alternative method of space transport is launching payloads into Earth's orbit using air-rocket (air-assisted) systems. The concept of mixed space transport involves launching a space rocket with a payload from an aircraft or other reusable platform-carrier at a specific altitude above the Earth's surface. The air-rocket system enables the launch of small satellites while reducing ground infrastructure and costs. Such a method seems promising in the context of relatively cheap, mobile and responsive small payload launch systems. It is an interesting proposition and beneficial especially for countries without convenient conditions to build their own spaceport. The paper analyses the fundamental requirements for the design of rockets used in this type of system and performs preliminary calculations of a conceptual two- and three-stage rocket capable of lifting a payload of 10 kg into a zero-inclination orbit at an altitude of 500 km. The analyses carried out were based on available research reports in this area.

Effects of Sweep on the Flutter of Wraparound Fins

In the field of missile and rocket design, efficiency is a crucial aspect considering factors such as weight, storage, stealth, and cost. Wraparound fins (WAFs) have been used for many years to control the trajectory of missiles and rockets because of their compactness and effectiveness. However, WAFs are structurally more flexible than other types of fins, which may negatively impact their dynamic behavior and aeroelastic response. This can cause catastrophic instabilities, such as flutter. To mitigate these challenges, it is essential to consider both aerodynamic and structural models of WAFs during the design process. In this study, the flutter velocity and frequency of WAFs were investigated in relation to different structural design configurations and sweep angles, in order to understand and improve the aeroelastic performance of swept WAFs.

Engineering calculation and analysis of aerodynamic heating for hypersonic rocket sled

Aerodynamic heating and temperature distribution are one of the problems that must be considered in the design of hypersonic rocket sled. In order to evaluate the aerodynamic heating environment of hypersonic rocket sled, the typical monorail rocket sled with cone cylinder combination is taken as the research object, the aerodynamic heating calculation methods such as Modified Newton Theory, Reynolds Analogy Theory, Reference Enthalpy Method and Fourier’s Law of Heat Conduction are used, and the temperature variation law of rocket sled conical surface under Mach 5 and Mach 7 ballistic is calculated and compared with the measured data in the literature to verify the accuracy of the calculation program. The engineering calculation method can provide a reliable design basis for the rocket sled structure and thermal protection configuration. The analysis of the calculation results shows that when the speed of hypersonic rocket sled exceeds Mach 5, serious thermal heating problem cannot be ignored, and protective measures such as thermal protective materials and helium channel should be taken.

PERFORMANCE OF SMALL SIZE AEROSPIKE. ENGINE

In this paper we are showing how aerospike nozzle can play an important role in rocket INDUSTRY as compare to other nozzles. We are using New rocket designs and increasing the performance of the current rockets But, the aerospike nozzle concept that has been under development since the 1950s is yet to be utilized on a launching. Due to its ability to adjust the environment by altering the outer jet boundary, the aerospike nozzle has better performance compared to present day others nozzle. An aerospike nozzle is designed for 20 bar pressure ratio to improve the performance of the aerospike nozzle for various conditions, some important design parameters and their performances were studied for low temperature flow conditions.

Dynamics model and numerical simulation of single stage water rocket

The initial launch parameters of a water rocket have a significant impact on its speed, altitude and other flight performance. The dynamic model describing the mass, velocity, nozzle velocity and internal pressure of water rocket was established by the force analysis of its ascending stage. Numerical simulation was carried out on the movement process of the water rocket. The changes of the internal pressure and thrust of the rocket body, and the speed and height law of the rocket in the flight process under the conditions of the same initial internal pressure and different initial loading propellant liquid were studied respectively. The results show that, under the designed conditions, the change law of the thrust force is similar to the variation of the internal air pressure in the ascending stage, and the internal air pressure decreases to 45%~57% of the initial value at last. Setting the volume of the initial loaded liquid to 0.43 helps to increase the boost duration and improve the utilization of the initial pressure, in which case, the maximum velocity of the launch rocket is not affected, but will lead to a decrease in thrust size.The above results provide important technical support for the optimal design of single-stage or multistage water rocket, and have strong guiding significance for carrying out water rocket flight test.

DYNAMIC MODEL OF THERMAL PROCESSES IN A LIQUID FLOWING THROUGH A CHANNEL

The article presents a dynamic thermal process model for an incompressible fluid moving through a channel of arbitrary cross-sectional shape. The study focuses on expressing the non-steady law of conservation of internal energy of the fluid by considering the dependence of parameters on both spatial coordinates and time. The one-dimensional spatial formulation of the problem is simplified by disregarding any changes in flow parameters in the direction perpendicular to the channel axis. The resulting equation represents a non-stationary, non-homogeneous, quasi-linear hyperbolic partial differential equation, which can be solved with the method of characteristics in the case of constant coefficients. The solution consists of two distinct solutions that are glued along the line x/t=v, with isolines shown in the figure provided. The research is relevant for the design of heat exchangers, rocket and aircraft engines, and other heat engineering devices.

Applications of Optimization Techniques for Solid Rocket Design

Solid Rocket Motors (SRM) are rocket motors that use solid propellants (fuel/oxidizer). Optimization of the motor’s design is the process to alter the schematics of the engine for higher engine efficiency and cost minimization. Solid Rocket Motor optimization is one of the key topics of aerospace engineering research today. Conventional methods for optimization have fallen obsolete due to the exceedingly large number of design variable in modern rocket engine optimization. This paper will summarize and review some of the related research carried out in this field. Modern methods of optimization are usually designed to be multidisciplinary using complex computerized hyper-heuristic algorithms like Genetic Algorithms, sequential quadratic programming, particle swarm optimization. In conclusion, this paper could be a useful reference for those who would like to acquire a scan of the field of modern SRM optimization, learn the core methodologies for the optimization process and understand the logic of related methods.

Utilization of two‐equation model for analysis of flow over a moving needle in the presence of magnetic field

AbstractMany applications of the ocean and aerodynamic engineering, like the design of submarines, missiles, aircraft, rockets, and jet engines, involve flow over surfaces. Moreover, several experimental propulsion methods for seagoing ships, submarines and aerospace crafts are based on magnetohydrodynamics (MHD). This research explores the MHD boundary layer flow over a moving needle (shaped like a paraboloid of revolution). Governing equations are modelled in cylindrical coordinates under axial symmetry constraint, which is then simplified by employing order of magnitude analysis. In the absence of a magnetic field, these equations are self‐similar, and this analysis is already available in the literature. We have applied the local non‐similarity method to derive the boundary layer equations up to the second level of truncation to compute the solutions of non‐similar boundary layer equations. The built‐in MATLAB routine bvp5c is used to solve these equations numerically. This analysis will assist in predicting flow quantities that are important in designing the objects mentioned above.

Scheme Design and Simulation Analysis of Recoverable Rocket Based on Rotor

The reuse of rockets is one of the most important methods to reduce launch costs. Based on the design scheme of the “Roton” space helicopter, this paper proposes a recovery method for rockets - rotors, which can not only accurately control the landing posture but also have high fuel utilization. To recover the first stage of the rocket, this paper carries out the scheme design and sets up the landing deceleration action with the engine reverse thrust and rotor lift. The simulation result shows that the scheme allows the first stage of the rocket to land safely, which provides a theoretical reference for engineering development in the future.