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Chelaru T.-V.,Polytechnic University of Bucharest | Chelaru A.,INCAS National Institute for Aerospace Research Elie Carafoli
UPB Scientific Bulletin, Series D: Mechanical Engineering | Year: 2015

The paper presents a random calculus model for calculating the precision of guided flight during terminal phase and automatic landing of a reentry vehicle. The proposed method is based on canonical decomposition of the random inputs, which allows us to obtain directly the output dispersion of the coordinates of the vehicle from the dispersion of any kind of random input signal, which passes through the differential equations of motions by using a decomposition of input signal on pulsation domains (PD) and by integrating the differential equation system for each PD. The novelty of the paper results from the theoretical method of random functions theory, applied to solve the technical problem of precision for guided flight during terminal phase and automatic landing of a reentry vehicle.


Chelaru T.V.,Polytechnic University of Bucharest | Danaila S.,Polytechnic University of Bucharest | Chelaru A.,INCAS National Institute for Aerospace Research Elie Carafoli
Proceedings of the International Astronautical Congress, IAC | Year: 2014

The paper presents a random calculus model for precision of guided flight during terminal phase and automatic landing of PRIDE vehicle. The proposed method is based on canonical decomposition of the random inputs, allowing solving a class of problems, which can be easily implemented as calculus software. The method consists of integration the equations of motion in linear form during terminal phase of the descending trajectory, considering influence of the random constraints as aerodynamic asymmetry, thrust command error, or noise sensors, using the canonical decomposition of random input functions. The method allows obtaining directly the output dispersion of velocity and coordinates of the vehicle from dispersion of any kind of random input signal, which pass thru differential equations of motions using a decomposition of input signal on pulsation domains (PD) and integrate the differential equation system for each PD. This method is approximate, because the number of PD is limited. Theoretically, if we use an infinite number of PD we can obtain the exact solutions. The results obtained in test cases show that this method, applied to the differential equations with the random input, gives good results if a high number of frequencies is selected. Although the solution appears to be complicated, leading to a high number of equations (4 times the number of frequencies), due to its symmetry and generality character, is a convenient method for solving these categories of problems, the majority of them having no analytic solutions. Furthermore, because mechanical system function like low pass filter, we can choose a limited number of frequency, which ensure a good accuracy of the results. In terms of results obtained with the considered calculus model, we will evaluate the average of terminal trajectory and landing position and the dispersion around them. The novelty aspect results in its technical purpose that to finding solutions for a real problem, using an adequate model from random function class. Even if the model may be subject to improvements, the results obtained there are technically acceptable and useful. The model proposed is an alternative to other models class, which uses random generated numbers, and can be use for cross checking between this two model classes. The novelty of the paper result from the theoretical method, of random functions theory, applied to solve the technical problem of precision for guided flight during terminal phase and automatic landing of PRIDE vehicle. Copyright © 2014 by the International Astronautical Federation. All rights reserved.


Chelaru T.V.,Polytechnic University of Bucharest | Chelaru A.,INCAS National Institute for Aerospace Research Elie Carafoli
RAST 2013 - Proceedings of 6th International Conference on Recent Advances in Space Technologies | Year: 2013

The paper intend to develop a calculus model for an innovative Reaction Control System (RCS) using hybrid rocket engine technology. Our RCS uses several hybrid micro-thrusters with their thrust modulated by a separate control system. For RCS, each of the thrusters will be able to burn a few minutes and its thrust will be modulated within certain limits by controlling the oxidizer flow. In order to reduce size and weight of the RCS we will use a single oxidizer tank which will have as output a flow distributor. The basic idea is not to stop any of the engines during system's operation but to minimize their thrust reducing the oxidizer flow. This approach is avoiding the inconvenience of repeated stopping and starting of the engine, which can create reliability problems to the entire RCS. By creating thrust imbalance between various hybrid micro thrusters, one can create torques with which the attitude or the trajectory of the vehicle can be adjusted. In terms of calculation model developed, it starts from our theoretical and experimental studies, which aimed to build a computational model for hybrid rocket engine highlighting the scalability, stability and its controllability. These studies were presented in RAST 2011 and are based on our own experiments performed in Electromecanica Ploiesti. Based on this concept we achieve a calculation of the performances of the RCS and an evaluation in their size. Conclusions and any discussion will be focused on technological possibilities for achieving the system and possible areas of application for the RCS. © 2013 IEEE.


Chelaru T.-V.,Polytechnic University of Bucharest | Chelaru A.,INCAS National Institute for Aerospace Research Elie Carafoli | Enache V.,Science Electromecania Ploiesti SA
RAST 2015 - Proceedings of 7th International Conference on Recent Advances in Space Technologies | Year: 2015

This paper develop a calculus model based on dedicated experiment for an innovative Reaction Control System (RCS) using hybrid rocket engine technology. Our RCS uses several hybrid micro-thrusters with their thrust modulated by a separate control system. For RCS, each of the thrusters is able to burn a few minutes and its thrust is modulated within certain limits by controlling the oxidizer flow. These studies were presented in RAST 2013 and are based on our own experiments performed in Electromecanica Ploiesti. Based on this concept we achieve a calculation of the performances of the RCS and make a comparison between theoretical and experimental results. Conclusions and discussions will be focused on technological possibilities to improve RCS performance and possible areas of application of it. © 2015 IEEE.


Bogos S.,INCAS National Institute for Aerospace Research Elie Carafoli | Dumitrache A.,Institute of Mathematical Statistics and Applied Mathematics | Frunzulica F.,Institute of Mathematical Statistics and Applied Mathematics | Frunzulica F.,Polytechnic University of Bucharest
AIP Conference Proceedings | Year: 2015

This paper aims to achieve a practical study about the quality of the aerodynamic results provided by CFD simulation compared with the high level confidence of the experimental results. Numerical simulation were developed at low Reynolds Number, specific for a high lift airfoil. Some ideas and strategies have been issued for choosing a suitable turbulence model or a suitable low Reynolds airfoil for UAV's or a wind turbine blade. © 2015 AIP Publishing LLC.


Bobonea A.,Polytechnic University of Bucharest | Bobonea A.,INCAS National Institute for Aerospace Research Elie Carafoli
AIP Conference Proceedings | Year: 2012

Wind turbine growth in size and weight made it impossible to control turbines passively as they were controlled in the past. Current efforts focus on increasing their aerodynamic efficiency and operational range through active flow control methods. One of the main methods of active flow control is the usage of blowing devices with constant or pulsed jets. By adding stored high-momentum air through slots into the boundary layer, they overcome adverse pressure gradients and postpone separation. Pulsed blowing sends short pulses rather than a continuous jet of fluid into the boundary layer and has been found to be more effective. Through CFD simulations over a 2D wind turbine airfoil, this research highlights the impact of different slot geometries with constant/pulsed blowing, on the effectiveness of this active flow control technique. © 2012 American Institute of Physics.


Bogos S.,INCAS National Institute for Aerospace Research Elie Carafoli | Stroe I.,Polytechnic University of Bucharest
UPB Scientific Bulletin, Series D: Mechanical Engineering | Year: 2012

This paper aims to achieve a study for increasing the level of confidence in the results concerning the lateral-directional stability of a real aircraft in the design stage, using the results from a flying scale model. Similarity coherent criteria are proposed for the dimensions, mass, inertia and cinematic characteristics between the real aircraft and the scale mockup model. Comparison between the real aircraft and the scale model plane show the same values for the damping factor in "Dutch roll". Values factored with a constant are obtained for the time characteristics in "Dutch roll", "Roll" and "Spiral" modes.


Chelaru T.V.,Polytechnic University of Bucharest | Pana V.,Polytechnic University of Bucharest | Chelaru A.,INCAS National Institute for Aerospace Research Elie Carafoli
Applied Mechanics and Materials | Year: 2014

The purpose of this paper is to present some aspects regarding the computational model and technical solutions for multistage suborbital launcher for testing (SLT) used to test spatial equipment and scientific measurements. The computational model consists in numerical simulation of SLT evolution for different start conditions. The launcher model presented will be with six degrees of freedom (6DOF) and variable mass. The results analysed will be the flight parameters and ballistic performances. The discussions area will focus around the technical possibility to realize a small multi-stage launcher, by recycling military rocket motors. From technical point of view, the paper is focused on national project "Suborbital Launcher for Testing" (SLT), which is based on hybrid propulsion and control systems, obtained through an original design. Therefore, while classical suborbital sounding rockets are unguided and they use as propulsion solid fuel motor having an uncontrolled ballistic flight, SLT project is introducing a different approach, by proposing the creation of a guided suborbital launcher, which is basically a satellite launcher at a smaller scale, containing its main subsystems. This is why the project itself can be considered an intermediary step in the development of a wider range of launching systems based on hybrid propulsion technology, which may have a major impact in the future European launchers programs. SLT project, as it is shown in the title, has two major objectives: first, a short term objective, which consists in obtaining a suborbital launching system which will be able to go into service in a predictable period of time, and a long term objective that consists in the development and testing of some unconventional sub-systems which will be integrated later in the satellite launcher as a part of the European space program. This is why the technical content of the project must be carried out beyond the range of the existing suborbital vehicle programs towards the current technological necessities in the space field, especially the European one. © (2014) Trans Tech Publications, Switzerland.


Chelaru T.V.,Polytechnic University of Bucharest | Pana V.,Polytechnic University of Bucharest | Chelaru A.,INCAS National Institute for Aerospace Research Elie Carafoli
Applied Mechanics and Materials | Year: 2014

The aim of this paper is to present a design method for the guidance navigation and control system (GNC) of a suborbital launcher. In order to achieve a symmetric evolution in the vertical plane we start with the decoupled form of the equation of motion. Afterwards these equations are linearized and the extended stability and command matrices are constructed by adding some auxiliary equations. The linear control law is obtained and the control matrix containing the unknown coefficients is presented. The design of the control system is based on a modified gradient method. To illustrate the proposed method the synthesis of the control system in the specific case of the SLT ("Suborbital Launcher for Testing - SLT" financed thru "Programme for Research- Development-Innovation on Space Technology and Advanced Research - STAR") is presented. © (2014) Trans Tech Publications, Switzerland.


Chelaru T.-V.,Polytechnic University of Bucharest | Barbu C.,Military Technical Academy | Chelaru A.,INCAS National Institute for Aerospace Research Elie Carafoli
RAST 2015 - Proceedings of 7th International Conference on Recent Advances in Space Technologies | Year: 2015

The purpose of this paper is to present some aspects regarding the mathematical model and technical solutions for small multistage launcher for used for scientific measurements. The model consists in numerical simulation of launcher evolution for different start conditions. The launcher model presented will be with six degrees of freedom (6DOF) with dynamical translational equations written in quasi-velocity frame. Model includes aerodynamics terms in polynomial form, the thrust terms and Guidance Navigation and Control relations, all of this adapted to 6DOF model. The results analysed will be the flight parameters and ballistic performances. The discussions area will focus around the technical possibility to realize this small multi-stage launcher, by using hybrid engine technology for uppers stage, including for Reaction Control System. © 2015 IEEE.

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