Rocket In Order Research Articles

The DESIGN AND BUILD OF A ROCKET STATIC THRUST TEST BASED ON A RASPBERRY PI 4.0

Abstract: A flying vehicle, or guided missile that gets a boost or a rapid reaction to the propellant material from the engine output is called a rocket. The Indonesian Army really needs measuring tools to determine the ability of a rocket to be made. Therefore, in this study the author will design a temperature measuring device, thrust, and temperature on a rocket with a rocket load of 1 ton. The research method used is pure experiment. This study aims to find data on temperature, thrust, and temperature on the rocket in order to find out the specifications needed so that the rocket can fly as desired. The resulting data will later be sent to the raspberry pi to be processed in order to get the results which will later be printed on paper that will be used for a research on making rockets to help the Indonesian Army in developing defense equipment in the rocket field.

Preliminary Study of Launch and Orbit of a CubeSat Using a Modified VSB-30 Launcher Vehicle

During the last decade, the world faced the mass insertion of small satellites in the space technology scenario. Every year, the number of micro and nanosatellites launched increases and gets more attention from players in the space market. Despite the lack of a national launcher, the Brazilian Space Program is known for some successful development in the last century, including its space assets, such as a privileged launch site near the equator, a family of flight proven and reliable sounding rockets for suborbital flights and microgravity experiments and universities with established small satellites programs. Thereby, the present work proposes a modification of the Brazilian VSB-30 sounding rocket in order to allow the launch and insertion in low Earth orbit (LEO) of small satellites fulfilling the gap of a national launcher. It also presents a CubeSat orbital decay simulation and orbital insertion simulation with the modified rocket launched from the Alcântara Launch Center as a matter of verifying the potential of national missions using this modified launcher.

Gas Pressure Measurement On Rocket Chamber Based On Strain Gauge Sensor

The use of rocket technology is currently also widely used for research. The rocket itself is a flying vehicle that moves by getting a boost through the combustion reaction that occurs in the rocket. The rocket itself can launch the load to the desired destination. In the field of science and technology, continuous efforts are required to realize the development of this rocket in order to achieve independence and prosperity. In the field of science and technology, continuous efforts are required to realize the development of this rocket in order to achieve independence and prosperity. The implementation of the strain gauge sensor through the MPXV7002DP gas pressure sensor obtained the amount of gas pressure during the combustion of the propellant in the rocket chamber, where the gas pressure can be measured, processed through the Arduino Uno microcontroller. The system can measure the maximum pressure up to 380 kPa and then the system can check any changes in voltage level to prove if there is a leak in the rocket.

Aerodynamic Coefficients Prediction for 122 mm Rocket by Using Computational Fluid Dynamics

In this study, the Computational Fluid Dynamics code in SolidWorks was used to predict the aerodynamic coefficients of 122 mm rocket in order to get more accurate aerodynamic coefficients and derivatives for flight simulation. This paper presented the governing equations of fluid dynamics applied in SolidWorks and developed the computational model of the rocket. The results were compared against calculated data from an empirical method. The conducted analyses showed that the comparison results are nearly the same. However, the empirical method provided a faster way to obtain results.

Pre-flight qualification test procedure for nanosatellites using sounding rockets

This work presents an innovative procedure to test nanosatellites subsystems on board sounding rockets. The procedure allows the subsystems to save their telemetry data during the whole rocket flight phases (including at the rocket lift-off). The subsystems of the FloripaSat (1U CubeSat) engineering model have been tested on board a VSB-30 rocket in order to validate their design and integration process. A dedicated embedded system has been proposed to operate as an electronic interface between the nanosatellite subsystems and the rocket electronics. Also, a system to process telemetry data was implemented on the ground station. The test procedure validates data frame definition, communication protocol specification, hardware integration specification, among other functionalities. Three FloripaSat subsystems have been tested: Electrical Power Subsystem (EPS); Telemetry, Tracking and Command (TT&C); and On Board Data Handling (OBDH). Several nanosatellites’ functions have been evaluated, including: battery monitoring; inertial measurement unit; temperature measurements; radio transceiver; and beacon. The flight results were considerably different from the results obtained at the laboratory. This has allowed important modifications to be made on the design of the CubeSat flight model as the revision on the radio circuit on TT&C and the implementation of an external battery charger circuit for the EPS.

Análise das percepções de estudantes do Ensino médio sobre o filme:"O Céu de Outubro"

O presente artigo apresenta uma atividade realizada com estudantes do 3° Ano do Ensino Médio sobre o filme “O Céu de Outubro”. Esse filme retrata o empenho de estudantes de uma pequena cidade que se dedicam na construção de um foguete, a fim de serem selecionados para a feira de ciências. Na atividade, os estudantes analisam criticamente aspectos/contextos abordados no enredo do filme. O trabalho realizado teve como objetivo propor aos estudantes uma análise crítica dos contextos familiares, políticos, históricos, sociais e a atividade científica voltada para a feira de ciências apresentados no filme, utilizando a proposta apresentada no artigo de Eduardo Fleury Mortimer e de Phil Scott, que propõe uma ferramenta para analisar as intervenções e as intenções no discurso do professor em sala de aula. O método de análise forneceu subsídios para inferir que o bom andamento da discussão do filme depende das intervenções do professor, o qual medeia o discurso para a intenção e o foco desejado, explorando a visão dos estudantes sobre os contextos apontados no filme.

Characterization of cosmic rays and direction dependence in the Polar Region up to 88 km altitude

Aims : The sounding rocket experiment REM-RED was developed to operate on board the REXUS-17 rocket in order to measure the intensity of cosmic rays. The experiment was launched from the ESRANGE Space Center (68 °N, 21 °E) on the 17th of March 2015 at the beginning of the most intense geomagnetic storm within the preceding 10 years. The experiment provided the opportunity to measure the intensity of cosmic rays in the Polar Region up to an altitude of 88 km above sea level. Methods : The experiment employed Geiger-Muller (GM) counters oriented with their axes perpendicular to each other in order to measure the cosmic ray intensity during the flight of the rocket. This measurement setup allowed performing direction-sensitive measurements as well. During the ascent phase the rocket was spinning and hence stabilized along its longitudinal axis looking close to the zenith direction. This phase of the flight was used for studying the direction dependence of the charged particle component of the cosmic rays. Results : In comparison with earlier, similar rocket experiments performed with GM tubes at lower geomagnetic latitudes, significantly higher cosmic radiation flux was measured above 50 km. A non-isotropic behavior was found below 50 km and described in detail for the first time in the Polar Region. This behavior is in good agreement with the results of the TECHDOSE experiment that used the same type of GM tubes on board the BEXUS-14 stratospheric balloon.

Pulsating Heat pipe Only for Space (PHOS): results of the REXUS 18 sounding rocket campaign

Two Closed Loop Pulsating Heat Pipes (CLPHPs) are tested on board REXUS 18 sounding rocket in order to obtain data over a relatively long microgravity period (approximately 90 s). The CLPHPs are partially filled with FC-72 and have, respectively, an inner tube diameter larger (3 mm) and slightly smaller (1.6 mm) than the critical diameter evaluated in static Earth gravity conditions. On ground, the small diameter CLPHP effectively works as a Pulsating Heat Pipe (PHP): the characteristic slug and plug flow pattern forms inside the tube and the heat exchange is triggered by thermally driven self-sustained oscillations of the working fluid. On the other hand, the large diameter CLPHP works as a two- phase thermosyphon in vertical position and doesn't work in horizontal position: in this particular condition, the working fluid stratifies within the device as the surface tension force is no longer able to balance buoyancy. Then, the idea to test the CLPHPs in reduced gravity conditions: as the gravity reduces the buoyancy forces becomes less intense and it is possible to recreate the typical PHP flow pattern also for larger inner tube diameters. This allows to increase the heat transfer rate and, consequently, to decrease the overall thermal resistance. Even though it was not possible to experience low gravity conditions due to a failure in the yoyo de-spin system, the thermal response to the peculiar acceleration field (hyper-gravity) experienced on board are thoroughly described.

Team Jarts Rocket Design

The objective of the 2013 Wisconsin Space Grant Consortium collegiate rocket competition was to construct a rocket to achieve an apogee of 3,000 feet and recover the rocket safely in flyable condition The rocket was designed to reach an apogee of 3000 feet. The main feature of the rocket includes passively controlling the drag to control the altitude reached. This is done by initial design of the airframe and fins. The other feature is deploying a drogue parachute on the ascent of the rocket in order to stop the rocket at 3000 feet. The rocket was designed to achieve an apogee greater than 3000 feet so that the deployment of a drogue parachute can stop the rocket at 3000 feet. Fiberglass tubing makes up the rocket due to its high strength and durability. Extra support was designed into the rocket to ensure that zippering does not occur. The rocket design included resulting in stable rocket based on the center of pressure and center of gravity. The recovery system of the rocket involves a dual deployment system. A drogue parachute is to be deployed at 2900 feet to stop the rocket at 3000 feet and control the rockets descent until the main parachute is deployed. The main parachute is then deployed at 700 feet and controls the rocket descent until landing on the ground. There is also a second set of ejection charges that will deploy the drogue parachute at apogee and the main parachute at 500 feet if the main charges do not deploy them. These ejection charges are fired by the MARSA4 and PerfectFlite StratoLogger altimeters. There is also a motor ejection charge backup based on a time delay to ensure the deployment of a parachute.Â

Simulación de la trayectoria de un cohete de dos etapas para posicionamiento de un nanosatélite en órbita

En el presente artículo se estudia el posible lanzamiento de un cohete de dos etapas impulsado por propelentes líquidos tipo RP-1 para transportar una carga útil de 4 kg equivalente a la masa de un nano-satélite de tres unidades de Cubesat. Se plantea el estudio con el diseño conceptual del cohete portador implementando un código computacional de optimización con la finalidad de reducir la masa al despegue y aumentar la velocidad. Se eligió como sitio de lanzamiento desde Colombia a la base del Grupo Aéreo del Oriente (GAORI) en Marandúa, Vichada, dado que en la actualidad se realizan investigaciones en lanzamiento de cohetes y ha sido utilizado por universidades para el desarrollo de tales investigaciones. Los valores obtenidos anteriormente son introducidos en un código computacional en FORTRAN desarrollado en el Observatorio Astronómico de la Universidad Nacional de Colombia para obtener el comportamiento de la trayectoria de vuelo del cohete durante su fase de ascenso e inyección orbital.

Acoustic measurements of the Epsilon rocket at liftoff

Launch vehicles generate intense acoustic field caused by the enormous thrust power during liftoff, and this acoustic field leads to harmful payload vibration. Japan Aerospace Exploration Agency (JAXA) plans to launch a new solid propellant rocket, Epsilon, in 2013. This three-stage rocket utilizes a reliable first stage motor of the JAXA’s H-2 rocket, SRB-A. Since the SRB-A is expected to cause excessive acoustic load, a countermeasure was required to mitigate the acoustic feedback to the vehicle. Computational works proposed a launch pad structure to attenuate the Mach wave radiation from the plume, and the acoustic wave generated by the plume impinging to the flame deflector. In the previous ASA conference, the authors introduced scale model tests using 1:42 scale rocket motors and the launch pad models. The scale model tests clarified the acoustic benefit of the launch pad structure. The tested geometry of the launch pad structure was adopted to the full-scale structure. Acoustic measurements are planned in the first launch of the Epsilon rocket in order to evaluate the acoustic influence with this newly constructed launch pad structure. The measurement setup and a brief review of the measurements (if possible) are discussed.

Trajectory Prediction for a Typical Fin Stabilized Artillery Rocket

This paper investigates the trajectory prediction and dispersion for unguided fin stabilized artillery rocket in order to explain the importance of the rocket production accuracy and the benefit of using guided rockets. The total dispersion results mainly from three effects. The first is the dispersion due to rocket production inaccuracy, which includes propellant mass, composition inaccuracy, rocket total mass, axial and lateral moments of inertia and resultant center of gravity. The second dispersion during boosting phase which includes launcher deflection, missile tip-off from the launcher, thrust and fin misalignments, and atmospheric disturbances such as tail wind, cross wind, and gusts. While the third is thedispersion during free-flight phase that is due to the fluctuations in wind profile. In this study, a trajectory calculation using a 6-DOF model was developed and applied for a typical artillery rocket, the 122 mm artillery rocket, at different mass and flight properties to predict the trajectory parameters and dispersion.

Effect of Variable Selection on Multidisciplinary Design Optimization: a Flight Vehicle Example

AbstractDifferent multidisciplinary design optimization (MDO) problems are formulated and compared. Two MDO formulations are applied to a sounding rocket in order to optimize the performance of the rocket. In the MDO of the referred vehicle, three disciplines have been considered, which are trajectory, propulsion and aerodynamics. A special design structure matrix is developed to assist data exchange between disciplines. This design process uses response surface method (RSM) for multidisciplinary optimization of the rocket. The RSM is applied to the design in two categories: the propulsion model and the system level. In the propulsion model, RSM determines an approximate mathematical model of the engine output parameters as a function of design variables. In the system level, RSM fits a surface of objective function versus design variables. In the first MDO problem formulation, two design variables are selected to form propulsion discipline. In the second one, three new design variables from geometry are added and finally, an optimization method is applied to the response surface in the system level in order to find the best result. Application of the first developed multidisciplinary design optimization procedure increased accessible altitude (performance index) of the referred sounding rocket by twenty five percents and the second one twenty nine.

Influence of range extension on dynamic stability for artillery rockets with wrap around fins

This paper presents the effects of the range extension of an artillery rocket with wrap around fins on its dynamic stability. The increased total impulse of the rocket engine for 50% and flight through the atmosphere layers with low air density extend the range for more than 75% of the nominal range. However, the flight of the unguided rocket through the rarefied air produces difficulties related to the dynamic stability. The small changes in the side moment which is induced by the angle of attack and Magnus moment of the improved rocket can produce instabilities in flight and increasement the angle of attack. The tolerances of the side moment were determined for both basic and improved type of the rocket in order to prevent instable flight. The influence of the range extension on the probable errors in range and cross-range was analyzed as well.

Development of C/C composites for OREX (Orbital Reentry Experimental vehicle) nose cap

The development of multiple-use type winged space shuttle planes (HOPE: H-II Orbiting Plane) is underway by NASDA (the National Aerospace Development Agency) and NAL (the National Aerospace Laboratories). In developing the HOPE, an Orbital Reentry Experimental vehicle (OREX) was mounted on the Number 1 H-II Rocket in order to collect basic data such as aerodynamic heat upon reentry to atmosphere. Carbon/carbon composite materials that have heat resistance of 1,600° C were used for a nose cap located at the nose tip part with the intention of using them in the HOPE. Because C/C composites are superior in specific strength and specific moduli with high heat resistance, they were adopted in the US Spaceshuttles. We were charged with manufacturing a nose cap made of C/C composites with a diameter of 1.7 m which is in the world's largest class. The OREX was launched from the Tanegashima Space Center in February, 1995, and the expected performance was attained, thus verifying the characteristics of a nose cap m...

Hardware-In-The-Loop Simulation for TR-I Rocket Roll Control System

A test vehicle, designated as TR-I rocket will be launched for acquiring design data and confirming the Solid Rocket Boosters (SRB) separation method for the H-II launch vehicle which has been under development by National Space Development Agency of Japan (NASDA).A simple roll control system has been devised for the TR-I rocket in order to satisfy complete linkage of the telemetry system and the requirements for the roll rate during the separation phase of SRB.In this paper, the roll control system and its Hardware-in-the-loop (HWIL) simulation test will be described.

Rocket measurements ofSqcurrents at midlatitude

Rubidium-vapor magnetometers were flown to altitudes near 150 km over Wallops Island, Virginia, in four Nike-Apache rockets. Two flights were made at times when no ionospheric current was expected, and none was encountered. The other two flights measured a current sheet with lower edge at 105 km and upper edge at 123 km, the total current being in agreement with that expected from the ground observations. One of the flights not encountering measurable ionospheric current was flown in conjunction with another rocket in order to obtain nearly simultaneous measurements of magnetic field, electron density, and winds. The ionospheric current calculated using the electron density and wind data was too small to have been observed by the rocket magnetometer.