Abstract: Sounding rockets are aerospace vehicles that were developed in the mid-20th century, and since then numerous investigations have been executed with the aim of innovate in this type of technology. However, the costs associated to the production of this type of technology are usually quite high, and therefore the challenge that exists today is to be able to reduce them. In this way, the main objective of this document is to present the design process of a Colombian aerospace mission capable to reach the thermosphere using low-cost “Candy” type solid fuel engines. This mission is the latest development of the Uniandes Aerospace Project (PUA for its Spanish acronym), which is an undergraduate and postgraduate research group at Universidad de los Andes (Bogotá, Colombia), dedicated to incurring in this type of technology. In this way, the investigations that have been carried out on Candy-type solid fuel, which is a compound of potassium nitrate and sorbitol, have allowed the production of engines powerful enough to reach space, and which represents a unique technological advance in Latin America and an important development in experimental rocketry. In this way, following the engineering iterative design methodology was possible to design a 2-stage sounding rocket with 1 solid fuel engine in each one, which was then simulated in RockSim V9.0 software and reached an apogee of approximately 150 km above sea level. Similarly, a speed equal to 5 Mach was obtained, which after performing a finite element analysis, it was shown that the rocket is strong enough to be able to withstand such speeds. Under these premises, it was demonstrated that it is possible to build a high-power aerospace mission at low cost, using Candy-type solid fuel engines. For this reason, the feasibility of carrying out similar missions clearly depends on the ability to replicate the engines in the best way, since as mentioned above, the design of the rocket is adequate to reach supersonic speeds and reach space. Consequently, with a team of at least 3 members, the mission can be obtained in less than 3 months. Therefore, when publishing this project, it is intended to be a reference for future research in this field and benefit the industry.
Abstract: Propellants based on Hydroxyl Terminated Polybutadiene/Ammonium Perchlorate (HTPB/AP) are the most commonly used in most of the rocket engines used by the Brazilian Armed Forces. This work aimed at the possibility of extending its useful life (currently in 10 years) by performing kinetic-chemical analyzes of its energetic material via Differential Scanning Calorimetry (DSC) and also performing computer simulation of aging process using the software Large-scale Atomic/Molecular Massively Parallel Simulator (LAMMPS). Thermal analysis via DSC was performed in triplicates and in three heating ratios (5 ºC, 10 ºC, and 15 ºC) of rocket motor with 11 years shelf-life, using the Arrhenius equation to obtain its activation energy, using Ozawa and Kissinger kinetic methods, allowing comparison with manufacturing period data (standard motor). In addition, the kinetic parameters of internal pressure of the combustion chamber in 08 rocket engines with 11 years of shelf-life were also acquired, for comparison purposes with the engine start-up data.
Abstract: In this paper numerical studies have been carried out
to examine the pre-ignition flow features of high-performance solid
propellant rocket motors with two different port geometries but with
same propellant loading density. Numerical computations have been
carried out using a validated 3D, unsteady, 2nd-order implicit, SST k-
ω turbulence model. In the numerical study, a fully implicit finite
volume scheme of the compressible, Reynolds-Averaged, Navier-
Stokes equations is employed. We have observed from the numerical
results that in solid rocket motors with highly loaded propellants
having divergent port geometry the hot igniter gases can create preignition
pressure oscillations leading to thrust oscillations due to the
flow unsteadiness and recirculation. We have also observed that the
igniter temperature fluctuations are diminished rapidly thereby
reaching the steady state value faster in the case of solid propellant
rocket motors with convergent port than the divergent port
irrespective of the igniter total pressure. We have concluded that the
prudent selection of the port geometry, without altering the propellant
loading density, for damping the total temperature fluctuations within
the motor is a meaningful objective for the suppression and control of
instability and/or thrust oscillations often observed in solid propellant
rocket motors with non-uniform port geometry.
Abstract: In order to determine the performance and key design parameters of rocket, the erosion of nozzle throat during solid rocket motor burning have to be calculated. This study aims to predict the nozzle throat erosion in solid rocket motors according to the thrust profile of motor in operating conditions and develop a model for optimum performance of rocket. We investigate the throat radius change in the static test programs. The standard method and thrust coefficient are used for adjusting into the ideal performance for conical nozzles. Pressure and thrust data acquired from the tests are analyzed to determine the instantaneous nozzle throat diameter variation throughout the test duration. The result shows good agreement of calculated correlation comparing with measured erosion rate data showing agreement within 1.6 mm/s. Nozzle thrust coefficient loss is found approximately 24% form nozzle throat erosion during burning.
Abstract: The ongoing effort to develop an in-house
compressible solver with multi-disciplinary physics is presented in
this paper. Basic compressible solver combined with IBM technique
provides us an effective numerical tool able to tackle the physics
phenomena and especially physic phenomena involved in Solid
Rocket Motors (SRMs). Main principles are introduced step by step
describing its implementation. This paper sheds light on the whole
potentiality of our proposed numerical model and we strongly believe
a way to introduce multi-physics mechanisms strongly coupled is
opened to ablation in nozzle, fluid/structure interaction and burning
propellant surface with time.
Abstract: In this paper parametric analytical studies have been carried out to examine the intrinsic flow physics pertaining to the liftoff time of solid propellant rockets. Idealized inert simulators of solid rockets are selected for numerical studies to examining the preignition chamber dynamics. Detailed diagnostic investigations have been carried out using an unsteady two-dimensional k-omega turbulence model. We conjectured from the numerical results that the altered variations of the igniter jet impingement angle, turbulence level, time and location of the first ignition, flame spread characteristics, the overall chamber dynamics including the boundary layer growth history are having bearing on the time for nozzle flow chocking for establishing the required thrust for the rocket liftoff. We concluded that the altered flow choking time of strap-on motors with the pre-determined identical ignition time at the lift off phase will lead to the malfunctioning of the rocket. We also concluded that, in the light of the space debris, an error in predicting the liftoff time can lead to an unfavorable launch window amounts the satellite injection errors and/or the mission failures.