Rocket Motor Thrust Research Articles

Data-driven deep learning approach for thrust prediction of solid rocket motors

Thrust is a crucial performance parameter of solid rocket motor. Accurately predicting the thrust is very important for rocket engineering monitoring and design. Traditionally, the study of solid rocket motor thrust has relied on model-based methods, which involve building complex models and requiring specialized knowledge. To address this challenging problem, this paper proposes a data-driven approach using deep learning to predict the thrust of solid rocket motor. The method utilizes raw data obtained from ground thrust tests and undergoes preprocessing steps, including segment-wise down sampling and sliding window sample preparation, to extract data features and enhance prediction accuracy. The approach employs a deep neural network architecture based on ConvNeXt to extract and learn advanced features from the preprocessed thrust data, resulting in an end-to-end framework for predicting solid rocket motor thrust. A significant number of experiments conducted on field data from solid rocket motors have confirmed the validity and superiority of the proposed method. The results indicate that both the proposed data processing and prediction methods perform well, with MSE, MAE, and MAPE values of 0.0047, 0.0170, and 1.1283 % on Dataset NO.1, respectively. On Dataset NO.2, the corresponding values are 0.0032, 0.0141, and 0.9846 %. Additionally, comparisons with other commonly used methods demonstrate that the proposed method generally exhibits lower errors. The experiments also investigated the key parameters that affect the performance of the model. The method provides robust support for rocket engineering monitoring and design, showing promising prospects in the field.

Research on Multi-axis Synchronous Test Technology based on GLINK Bus

As a rocket engine for thrust vector control system of high reliability and redundancy, high precision, and the requirements of lightweight, servo system for test system can quickly test a certain level and fast data processing of the test location, transient, and frequency characteristic indexes, a rocket motor thrust vector control system using GLINK as control bus. Moreover, the first level servo is an all-in-one machine with four drives and controls, so the test system is required to have the ability to control and test four actuators simultaneously under the GLINK main line. This paper discusses the hardware knowledge of the test system based on the GLINK bus: GLINK bus technology, communication topology of GLINK station card and network card; software function of the test system; multi-axis synchronous test function, and synchronous data processing function developed by LabVIEW.

Cooling of rocket plume using aqueous jets during launching

With an increase in rocket motor thrust, the thermal loads on the launch platform by the impingement exhaust jet are significantly increased, this poses serious safety risks to rocket launch. Therefore, a water cooling system should be designed to improve the thermal environment. In this study, the cooling effects of a water spray system with different injection velocities were analyzed through numerical simulation of high precision, and the afterburning effect of the exhaust gas was also considered to improve the calculation accuracy. The results showed that afterburning primarily occurred in the mixing layer, leading to a larger high-temperature region in the impingement plate. The water cooling system realized thermal protection for the launch platform, which was achieved by decreasing the temperature and velocity of impingement jet. With an increase in the aqueous jet velocity, a more significant cooling effect of the water spray system was obtained. However, the guiding performance of the deflector was decreased owing to the large amount of water vapor formed in the channel. It is suggested that the water spray system design should adopt an aqueous jet velocity of 20 m/s as a parameter after considering the cooling effect and guiding performance of the deflector.

Variability analysis of ABS solid fuel manufactured by fused deposition modeling for hybrid rocket motors

The process of fused deposition material (FDM) was used to manufacture propellant grains of Acrylonitrile Butadiene Styrene (ABS) as novel rocket fuel grain, with three types of geometry in the burning port. These solid fuel grains were used to measure the typical characteristics of combustion in rocket motors such as thrust and pressure inside the combustion chamber, seeking to obtain preliminary characteristics of operation and analyze the effect of combustion port geometry on pressure and thrust, using Multivariate Analysis of Variance (MANOVA) as statistical method. Two of the three geometries were manufactured with a helical-finocyl configuration, specially designed to be fabricated by Direct Digital Manufacturing (DDM), the other one was a straight-bore geometry also by DDM. This characterization experiment was performed on a static hybrid rocket engine, designed to inject 99.98% pure nitrous oxide into a combustion chamber with capacity to withstand 6.9 MPa of pressure, with an easy-to-exchange nozzle, avoiding erosive behavior in the throat. Statistical analyses made with the ABS fuel grains, suggest a significant effect on rocket motor pressure and thrust, due to helical geometric changes made to the combustion port of solid fuel grains made by FDM manufacture process.

Numerical Evaluation of the Effects of Inclusions on Solid Rocket Motor Performance

This work describes the application of original computer graphics methods to examine the effect of inclusions within the solid rocket motor thrust chamber. The adoptance of a dynamic three-dimensional triangular mesh with a self-intersection removal algorithm, as the core of the burning surface regression module, allows the aforementioned target to be reached. Many phenomena not measurable in an analytical closed form, like different-shaped voids within the grain, can be investigated. Indeed, both local heterogeneities and surface imperfections could generate unexpected variations in the combustion chamber pressure pattern and early thermal protections exposure. The aforementioned techniques are developed and tested on a ballistic simulator developed using computer graphics methods to improve mesh handling. Simulation results are carried out and discussed.

Performance analysis for aerial-towing system

The dynamics of open chain systems have been widely investigated as a result of its increased applications such as aerial towing system, submerged towed cable applications, robot manipulators, and bio-mechanical applications. In this study, a practical engineering system, Mine Clearing Line Charge MICLIC system is taken as an example for an open chain system with accelerated end and varying number of segments. To validate the proposed model, a comparison between two different finite segment methods namely Kane’s method and Riccati transfer matrix method of linear multibody system (Riccati-MS-TMM) have been performed for a large open chain system. Results show that computational speed of Riccati-MS-TMM outperforms Kane’s method due to its huge matrix size to be solved. Then, using a predefined rocket motor total impulse, a parametric study has been performed on MICLIC system to illustrate the impact of different rocket motor thrust profiles on the resulted flight performance including the missile range, flight time, maximum tension applied on the rocket towing bridle, and the missile stability.

A 2-DOF Model of an Elastic Rocket Structure Under Circulatory Force

It is intended, in this paper, to develop a mathematical and numerical model of an elastic space rocket structure as a Beck’s column excited by a follower (or circulatory) force. This force represents the rocket motor thrust that should be always in the direction of the tangent to the structure deformed axis at the base of the vehicle. We present a simplified two degree of freedom rigid bars discrete model. Its system of two second order nonlinear ordinary differential equations of motion are derived via Lagrange’s energy method, allowing for a general understanding of the main characteristics of the problem. The proposed equations consider up to third order (cubic) inertia, stiffness and forcing terms. Among other rich nonlinear dynamic behaviour of this model, depending on parameters and initial conditions choices, either stable or unstable limit cycle post-critical steady state solutions are possible. The latter is a form of flutter. Numerical simulations are carried out using Runge-Kutta’s 4th order algorithm in Matlab environment.

Experimental and simulation testing of thermal loading in the jet tabs of a thrust vector control system

The paper discusses the temperature changes in mechanical jet tabs in a system of rocket motor thrust vector control, estimated by the simulation and experimental tests methodology. The heat transfer calculation is based on complex computational fluid dynamics simulations of both the nozzle and external tab flows, as the comprehensive integral flow zones with high flow parameters gradients. Due to a complexity of the model for flow calculations, the experimental estimation of the calculated results is carried out. The temperature is measured by jet tabs embedded thermocouples, and conducted through the rocket motor static tests. A good agreement of the calculated and measured results is achieved. The main aim of the developed method is to establish an approved calculation tool for designing new TVC systems in order to avoid disadvantages due to overheating.

New Capabilities and Discovered Interconnectivities for a Curriculum-Integrated Multicourse Model Rocketry Project

To provide students a more coherent and cohesive view of the mechanical engineering curriculum, we created and are delivering a multicourse curriculum-integrated engineering project that permeates and unifies five required classes within our undergraduate curriculum: 1) Freshman Design, 2) Dynamics, 3) Numerical Analysis, 4) Fluid Mechanics, and 5) Thermodynamics. Students enrolled in these Rocket Project (RP) classes design, build, flight test, and analyze model rockets through hands-on exercises. These activities challenge students to work on different aspects of the same rocket project across all four years of their degree program. Critical to the seamless collection and presentation of data and experimental/numerical techniques across five courses was the development of new laboratory, field, and simulation capabilities driven by our goal: to measure all unknown variables needed for rocket performance analysis and modeling in-house without reliance on external data. These needed capabilities included: 1) collecting acceleration and barometric altitude data from a model rocket flight, 2) simulating via computer rocket trajectories for comparison to actual measured altitudes, 3) evaluating rocket performance by numerical methods to validate modeling assumptions, 4) determining rocket drag coefficient as a function of Reynolds number for velocities relevant to a launch, and 5) measuring rocket motor thrust as a function of time as well as the energy density of the fuel used. As these capabilities were developed, additional course interconnectivities and opportunities for data sharing were discovered and exploited to further enrich the course experience for students.Â

Orbital Injection Errors and Sensor Requirements for NERVA Space Launchers

The input module of the NERVA space launcher guidance system consisting of the inertial and sensor platform is responsible for the basic accuracy if the ascent trajectory and injection efficiency. The sample rate magnitude and data filtering along the real time trajectory are the only tools available for improving the guidance accuracy up to the level of requirements to secure admissible orbital injection error and the subsequent flight corridor during the orbital ascent. Analysis of the NERVA-1 flight telemetry flow from the onboard inertial platform raises the problem of the optimal selection of the onboard sample rate and of the rate of telemetry, which are not identical. The orbit injection errors are chosen from the orbit altitude constraints and subsequent accuracy requirements for the inertial sensors are derived. They show that the accuracy requirements are moderate and may be covered with almost conventional sensors. To improve the flight guidance accuracy the rocket motor chamber pressure and thrust are measured and observation of the preflight zero drift, recording noise and of the high level of embedded noise during both powered and coast atmospheric flight is performed. Simple filtering based on frequency Fourier analysis is delivered with conclusions regarding the intelligent algorithm enhancement that are developed and implemented on the next generation of flight research drone missiles RT-759M NERVA-2, right in preparation. The main rationale of that algorithm stands in the method of discriminating between false and true information on each measuring point immediately after the data are delivered by the sensors. Learning procedure from previous preflight recordings and from gradual accumulation of concurrent data streams subjected to FF spectral analysis are combined to improve data filtering, for immediate release to the next module of the autopilot. The rate of sampling is optimized from the analysis of the previous flight, inertial data records and test stand pressure and thrust records that show the level of noise. The behavior of the electronics under the dynamical loads of the rocket flight, involving overloads of more than 20 g-s and the level of vibration during the real flight and other sources of measuring errors are also focused in the research. During simulated work of the sensor platform the algorithm has been acceptably validated and prepared for real flight test performance. Information important for the NERVA autopilot design activity is structured through the multiple variance approach.

Testing pyrocartridges and the rocket motor of the ejection seat

Due to a complex geometrical shape of the rocket motor of the ejection seat, the rocket motor thrust occurs under certain angle in relation to the plane of symmetry of the ejection seat. A number of tests were carried out in order to determine thrust intensity and angle of attack. The rocket motor thrust was measured on the multicomponent test stand. The ejection seat whit a dummy simulating a mass of a pilot was tested during ejection. The paper presents an analysis of the ejection seat flight in the initial phase of ejection, during the rocket motor running. The results of the acceleration read-outs were processed and then used for the determination of the characteristics of the ejection seat flight. A theoretical model of the ejection seat flight is given in the paper. .

Uncertainty analysis of rocket motor thrust measurements with correlated biases

The uncertainty analysis of a rocket motor thrust measurement and calibration system is presented. The thrust measurement system has 10 load cells with 5 in tension and 5 in compression, and the calibration system has 6 load cells. The elemental sources of load-cell random uncertainty considered are excitation voltage, amplifier stability and noise, and analog-to-digital conversion resolution. The elemental sources of load-cell systematic uncertainty (bias) considered are calibration, excitation voltage, nonlinearity, hysteresis, system alignment, and other mechanical sources such as friction. The calibration hierarchy and excitation systematic uncertainties are correlated since they have a common source. These correlation effects dominate the resulting thrust measurement uncertainties and form the focus of this study. The trust measurement uncertainty is up to three and a half times larger when correlation effects are considered relative to the case where correlation is neglected.

High-performance missile synthesis with trajectory and propulsion system optimization

Synthesis of a high-performance two-pulse-motor-propelled missile by simultaneous optimization of trajectory and propulsion systems in two operational scenarios is described. This work employed a quasiNewton parameter optimization scheme with penalty functions to meet terminal and path constraints. The trajectory control variables were parameterized using piecewise linear open-loop commands and piecewise constant linear feedback gains. The pulse motor parameters optimized were pulse split, average thrust levels, neutrality factors, pulse burn times, and the interpiilse delay. Optimization of the rocket motor thrust time curve in addition to the trajectory increased the range by up to 11% when compared with the trajectories optimized with originally specified rocket motor propulsion data.

Mechanical noise suppressor for small rocket motors

A noise suppressor for small rocket motors including a plurality of perfoed metal cylinders disposed in concentric relation and secured to a support plate. A noise suppression material is disposed in the chamber of each adjacent cylinder. Noise suppression material is also disposed in the center cylinder and is expelled by the rocket motor thrust. A collar on the support plate secures the motor to the suppressor.