Satellite Launch Vehicle Research Articles

Theoretical Investigations to Random Testing Variants and its Implications

In real world Testing is the most challenging job in software development. In these days of the information technology epoch, the software has become the lifeline of every human activity. It may not be hyperbole, if someone states that our lives will come to stand-still if every software machine in the universe stops working! When software plays such a crucial role in our lives, it is very important that the software we use should be very high quality and reliability. Even the small gremlin is there in any software that causes huge disaster. These are the examples PATRIOT MISSILE- the patriot missile example shows how a small software bug accumulation of error due to rounding off a number can lead to catastrophe. And ARIANE-5- European Space Agency launched the ariane5 satellite launch vehicle. Exactly 40 seconds after lift-off at an attitude of 3700meters, the launcher devastated and became a ball of fire. This type of mint errors can’t be covered in normal testing’s like Unit Testing, Integration Testing, etc. For that reason RANDOM TESTING is introduced. Random testing is a dominant or a commanding tool in finding low-frequency bugs that are nearly impossible to find using other methods. A Low-frequency bug sometimes leads to system to crash, hence these have to be identified and removed. By using a test, Oracle one may also detect non-crashing errors in the system. Here the objective of work is to study the advantage and various implications of Random Testing, later we will develop one assumed application “Trading system”, to that we will apply random testing to accentuate theoretical and practical implications including variants of Random Testing.

Metallurgical Analysis of a Failed 17-4 PH Stainless Steel Pyro Bolt Used in Launch Vehicle Separation Systems

Abstract Precipitation hardenable stainless steels are widely used in aerospace applications because of their excellent combination of strength and corrosion resistance. In view of this advantage, Fe-17Cr-4Ni (17-4 PH) steel is used in pyro bolts in space ordnance systems for stage separation in a satellite launch vehicle. These systems typically consist of a pyro charge to sever a given material resulting in stage separation, and these materials are exposed to high strain rates. In one such application, although perfect severance took place, secondary cracks were noticed longitudinally, which are not desirable. This steel, in view of its composition, has delta ferrite in a martensitic matrix in the hardened and tempered condition. In wrought products worked unidirectionally, delta ferrite gets oriented along the direction of thermomechanical processing. The effect of delta ferrite orientation on the impact properties of 17-4 PH stainless steel was studied in various heat-treatment conditions with a view to understand the effect of microstructure on impact properties. It was found that the transverse impact properties are severely affected by the presence of highly oriented delta ferrite stringers for any given heat-treatment condition. This has been attributed to the weak delta ferrite–martensite interface, which facilitates easy crack propagation. Finally, it was suggested that the anisotropy in impact properties need to be carefully considered in the design of components for applications that involve high strain rates.

Infrasonic propagation modeling of orbital launch vehicles

Over the last two years, NASA’s Wallops Flight Facility has been launching medium lift rockets for experimental and space station resupply missions. These launches have been a great opportunity to use the rocket-generated infrasound as a repeatable source to study the long range propagation over different seasons. For some of the launches during this period, two, 20-m arrays were deployed along different directions from the launch site. To increase the number of available sensors for comparison to three-dimensional propagation model results, data from the US Array was incorporated into the study. The US Array data significantly increased the range and azimuths used to compare the propagation modeling and measurement results. Calculated and measured transmission losses will be discussed.

Indian space transportation programme: Near term outlook and issues for commercialisation

The Indian space transportation programme has grown from strength to strength with the launching of sounding rockets in the 60׳s to the development of heavy lift vehicles for telecommunication satellites in the present decade. With the growing market confidence in Indian Space Research Organisation׳s ability to reliably deliver payloads to low Earth orbit with its Polar Satellite Launch Vehicle, there is an inherent opportunity for India to cater to the commercial market. The present work assesses the current launch capacity of India in retrospect of international launches and provides India׳s outlook for the space transportation in the current decade. Launch capacity correlation with the requirements within the Indian space programme as well as the current space transportation infrastructure have been considered to identify bottlenecks in catering to the current national requirements alongside securing a greater market share in the international launch market. The state of commercialisation of launch vehicle development has been presented to provide an overview of policy and organisational issues for commercialisation of space transportation in India.

Generalized Dynamic Inversion Scheme for Satellite Launch Vehicle Attitude Control

This paper presents a generalized dynamic inversion attitude control design for satellite launch vehicle. Dynamic constraints that encapsulate the attitude tracking objectives are prescribed in the form of asymptotically stable differential equations in the attitude variables deviations from their reference trajectories. The constraint differential equations are evaluated along the vehicle's attitude trajectories and are inverted to solve for the control variables using the Greville formula for underdetermined algebraic systems. The associated null control vector is constructed using Lyapunov stability principle, and a delaying dynamic scaling factor is augmented in the involved Moore-Penrose generalized inverse for assured singularity avoidance and stable attitude tracking. The effectiveness of the proposed control scheme is shown by its evaluation on a 6DOF model of a satellite launch vehicle in the presence of various external disturbances.

Dr. A.P.J Abdul Kalam-"A Man Beyond Science".

Dear Sir, The extraordinary charismatic personality of Dr. Kalam has sensitized every individual in our nation. His reach to the people was beyond the bar of caste, color and creed. An ordinary man, who came from a humble family of Rameswaram, Tamil Nadu, India, reached to the highest national rank by being the 11th President in Indian Republic. He has been a very sincere worker of his field and was a continuous source of inspiration for all generation of people. Among the awards which he deservingly earned in his lifetime, few of them include Padambhushan (1981), Padma Vibushan (1990), Bharat Ratna (1997) King Charles Medal-II (2007) and many more. However, the greatest award was the respect he commanded among common man and specially youth for whom he is an unyielding source of inspiration. His speeches and his perspective towards “Science and its Utilization for People” are a worth beyond words. He graduated from Madras Institute of Technology in aerospace engineering and later remained associated with Defense Research and Development Organization for nearly 40 years. He started his career with a project of developing a small helicopter for Indian army. Dr Kalam developed India’s first indigenous Satellite Launch Vehicle (SLV-3) which successfully placed the satellite, Rohini in the near orbit of earth in July 1980 and made India an exclusive member of Space Club. Subsequently, he played important role in development of missiles like Agni and Prithvi which made India a self sufficient in its Defense Power. He was called as “Missile Man” for his contribution in missiles development. Dr Kalam has played a pivotal role as Chief scientific Advisor to Prime Minister between July 1992 to December 1999, it was during this time that “Pokhran-II” the nuclear test was conducted. His hard work in this field made India a Nuclear Power. Dr Kalam had significant contributions to medical science which are less discussed in Indian media and literature. He advocated, the role of biotechnology in medical science research, emphasized the use of science and technology for the benefit of common man. India’s first indigenous low cost cardiac stent was developed by him in collaboration with Dr. Somaraju Bhupathiraju known as Kalam-Raju stent during 1994 – 1996. It reduced the cost of stents by 1/4th compared to prevailing market rate. Approximately 2000 Kalam-Raju cardiac stent has been fitted to many needy patients and showed the path of affordable health care [1]. Although, the medical field has progressed to drug eluting stents, this invention was a landmark in its time. In 2012, they developed a tablet computer for purpose of rural health care service, which was named the ‘Kalam-Raju tablet’. One of his great achievement was the development of light weight calipers for polio affected children by using carbon-carbon and carbon-polymer materials (glass-filled polypropylene which are used in missiles) during 1995 – 1996, thus reducing weight of the caliper to 1/10th of the original weight. Over 50,000 children were benefitted with these calipers which reduced their pain significantly, while walking. Throughout his life, he took up the mission to ignite and inspire young minds towards national spirit by meeting school students across the country. He has also written many thought provoking and inspirational books such as “ Wings of Fire”, “ Ignited Minds” and “ India 2020”, which has become house hold names with youth of India. During the last decade, Dr. Kalam has addressed over five million youth below the age of 17 and inspired them to become an active participant of India Vision 2020. He had addressed several Children Science Congresses across the country. Though, everyone who is born has to leave for the heavenly abode, but Dr. Kalam has left the legacy of his scientific achievements, thoughts and teachings as a source of inspiration for generation to come in future India and the world. Today, we really are gratefull to the almighty for giving us the opportunity to live in an era when we could share the sunlight with such a great personality. May his soul rest in peace.

Reconfigurable Direct Allocation for Multiple Actuator Failures

Flight control systems are often equipped with reconfigurable control allocation schemes to redistribute control commands in the event of actuator failures. In this brief, a reconfiguration scheme is proposed and is based on the direct allocation method. An iterative algorithm is developed and uses the same lookup tables prepared offline for normal operations, thereby saving a considerable amount of computation time and process memory. To elucidate the effectiveness of the method, the proposed allocation algorithm is applied to a satellite launch vehicle model. Simulation results show satisfactory performance of the method in the presence of multiple actuator failures.

Новое поколение инерциальных навигационных систем на основе ВТГ для аппаратов, обеспечивающих запуск спутников

Новое поколение инерциальных навигационных систем на основе ВТГ для аппаратов, обеспечивающих запуск спутников

Mathematical modelling of a cryogenic engine

Cryogenic propulsion systems using liquid hydrogen and liquid oxygen propellant combination are used in satellite launch vehicles in view of the higher specific impulse (ISP). Due to extreme low temperature of cryogenic propellant and low density and explosive nature of LH2, the development of cryogenic propulsion system is very complex and time consuming. In ISRO an indigenous cryogenic upper stage powered by an engine developing a nominal thrust of 73.5 kN in vacuum with a propellant loading of 12.8 tonnes is developed and successfully flight tested for the first time in GSLV D5 flight on 5th January 2014. A mathematical model is developed for finalizing the engine start and shut off sequence, ensuring smooth and safe ignition, predicting performance of subsystems under transient phase, nominal and off nominal conditions, control systems parameter settings and for the performance estimation of engine in sea level and flight. This paper highlights the configuration and working of a cryogenic engine, mathematical modelling of the engine, applications of the model in the cryogenic engine development and comparison of predicted values with the test results.

Orbit Injection Considerations for Cluster Flight of Nanosatellites

Cluster flight is defined as the operation of multiple cooperative satellites under minimum and maximum distance constraints. Unlike traditional orbit injection of multiple noncooperative satellites, judicious constraints should be imposed on the orbit injection of a cluster. Moreover, the orbital injection should enable a fuel-efficient cluster establishment maneuver. This study develops constraints for the orbit injection of a satellite cluster. Evaluation of feasible orbit injection scenarios is performed. Particular launch vehicles and a concomitant set of orbital injection scenarios are found, which are the most suitable for long-term operation of a cluster. In addition, an efficient cluster establishment control method is developed, which prevents the cluster from drifting apart too rapidly. Several orbit injection scenarios are simulated and their feasibility for a planned cluster flight mission is evaluated. It is shown that both the Polar Satellite Launch Vehicle and the Dnepr launcher can be used. In addition, a new customized release procedure is developed, which guarantees cross-track separation among the satellites, a desirable feature in a number of practical missions.

Lifetime Estimation of the Upper Stage of GSAT-14 in Geostationary Transfer Orbit.

The combination of atmospheric drag and lunar and solar perturbations in addition to Earth's oblateness influences the orbital lifetime of an upper stage in geostationary transfer orbit (GTO). These high eccentric orbits undergo fluctuations in both perturbations and velocity and are very sensitive to the initial conditions. The main objective of this paper is to predict the reentry time of the upper stage of the Indian geosynchronous satellite launch vehicle, GSLV-D5, which inserted the satellite GSAT-14 into a GTO on January 05, 2014, with mean perigee and apogee altitudes of 170 km and 35975 km. Four intervals of near linear variation of the mean apogee altitude observed were used in predicting the orbital lifetime. For these four intervals, optimal values of the initial osculating eccentricity and ballistic coefficient for matching the mean apogee altitudes were estimated with the response surface methodology using a genetic algorithm. It was found that the orbital lifetime from these four time spans was between 144 and 148 days.

Infrasonic energy from orbital launch vehicles

Large, heavy-lift rockets have significant acoustic and infrasonic energy that can often be detected from a considerable distance. These sounds, under certain environmental conditions, can propagate hundreds of kilometers from the launch location. Thus, ground-based infrasound arrays can be used to monitor the low frequencies emitted by these large rocket launches. Multiple launches and static engine tests have been successfully recorded over many years using small infrasound arrays at various distances from the launch location. Infrasonic measurements using a 20 m array and parabolic equation modeling of a recent launch of an Aries V rocket at Wallops Island, Virginia, will be discussed.

전개형 위성 안테나 메커니즘 모델링 및 분석

지구 관측용 위성의 한 종류인 SAR(Synthetic Aperture Radar) 위성은 지구의 영상 획득 시, 주변 환경의 영향을 받지 않는다는 장점이 있어 현재까지 많은 수가 개발되었다. 높은 영상 품질을 얻기 위해 SAR 안테나는 큰 직경을 가져야 한다. 그러나 위성을 우주까지 운반하는 발사체의 내부 공간은 한정적이므로 SAR 안테나는 발사체 내부 배치 시에는 접어놓았다가 우주에서는 펼쳐질 수 있는 전개형으로 설계된다. 본 연구에서는 위성의 목표 임무를 고려하여 전개형 위성 안테나의 메커니즘과 관련된 설계 요소 값들을 선정하였다. 이를 바탕으로 전개형 위성 안테나의 형상을 설계하였으며, 시뮬레이션을 통해 위성 기동 시 전개형 위성 안테나에 발생하는 변화를 관찰하였다. Large number of SAR(Synthetic Aperture Radar) satellites, one type of earth observation satellite, have been developed as they have the advantage of not being affected by surrounding environment during the earth image acquisition. In order to gain high image quality, SAR antenna should have large diameter. However, internal space of satellite launch vehicle is limited and this leads SAR antenna to be designed deployable so that it can be folded in launch vehicle and unfolded in space. In this research, values of various design factors of deployable satellite antenna were chosen considering satellite's target mission. Configuration of deployable satellite antenna was designed by applying the chosen values of design factors, and variation in deployable satellite antenna during satellite maneuver was observed through simulation.

Failure analysis of cracked brass turrets used in electronic circuits

Copper alloys are widely used in electrical circuits mainly due to their electrical conductivity. In electrical circuits, solder terminals are used to provide multiple electrical connections and swage type turret terminals are preferred owing to their high reliability under dynamic environments where the electrical systems are subjected to higher vibrations and line currents, as in the case of satellite launch vehicles. Owing to high machinability, free-cutting brass is preferred in the manufacture of such small sized terminals, which will be cold worked/swaged to the printed circuit boards. Improper selection of the composition of free-cutting brass can lead to cracking of the terminals, thereby affecting the reliability of the system. This paper describes cracking of turret terminals made of free cutting brass during swaging operation. The crack propagation tendency during cold working of brass turret terminals made from different compositions is studied through optical and scanning electron microscopy. Finally, composition for obtaining crack free terminals during swaging is suggested for applications involving cold swaging.

Comparison of High Cycle Fatigue in 4340 and 300M Steel Welded with Fiber Laser

The AISI/SAE 4340 steel and 300M have been used in severe conditions where high strength and toughness are necessary. Since there is a particular interest in the aerospace industry, they are being used in aircraft landing gear, rocket engines vessels and satellite launch vehicles. Therefore, the understanding of the weldability of these steels has been the subject of special interest. This study aims to evaluate and compare the microstructure and mechanical strength by tensile and fatigue in steel sheets of 4340 and 300M steels after laser welding and tempering heat treatments. The results showed that the welds had high hardness and small heat affected zones. Nevertheless, there was not a decrease in the uniaxial tensile properties for the 300M and 4340 steels. The decrease in fatigue life due to the welding process, compared to the unwelded condition, was small, indicating that the parameters welding were adequate and that this process is viable.

Parameterization of Micro Satellite Launch Vehicle

The new trends in micro satellites mission for wide range of applications has motivated worldwide research in design and development of micro satellite launch vehicles (MSLVs) as a vital options in the deployment of such satellites into their respective mission orbits. Presently, most micro satellites are launched through rideshares with rockets or rocket-like systems, which were developed essentially in the seventieth and eightieth of the last century. The MSLV as a proportional micro satellite launch vehicle, up till now has only the status of technology and system study work and is well known that the research and development of MSLV are diversified in different fields. However, the realization of MSLVs is rigorous since such vehicles compound the inherent difficulties of traditional launch experiment with intrinsic problems of uneven mass distribution, slenderness, high frequency and amplitude of vibration etc. MSLV are also more likely to be affected by wind gusts and other disturbances than their conventional counterpart, and the physical parameters such as mass distribution and moments of inertia can be easily altered by changing the payload location and unpredictable propellant burning. All these factors compromise the additional velocities to overcome the aerodynamic drag, to account for steering of the vehicle and other losses. Consequently, We prototype a three stage micro-satellite launch equipment as a similitude of the European Space Agency’s VEGA launcher using a scale down derivation model and parameterized using existing VEGA launch vehicle (LV) technical data. VEGA launch data, total orbital velocity requirement including the losses and stage payload ratio are inferred and applied to the model of parameterization of launch vehicles through coupledparameters simulations. This approach can be helpful in viability and performance insights in terms of payload mass, stage mass distribution and material properties precision and effectiveness. It can also serve as a background for the development of time optimal waypoint trajectory of micro satellite launchers.

Development of a Numerical Model for an Expanding Tube with Linear Explosive Using AUTODYN

Pyrotechnic devices have been employed in satellite launch vehicle missions, generally for the separation of structural subsystems such as stage and satellite separation. Expanding tubes are linear explosives enclosed by an oval steel tube and have been widely used for pyrotechnic joint separation systems. A numerical model is proposed for the prediction of the proper load of an expanding tube using a nonlinear dynamic analysis code, AUTODYN 2D and 3D. To compute a proper core load, numerical models of the open-ended steel tube and mild detonating tube encasing a high explosive were developed and compared with experimental results. 2D and 3D computational results showed good correlation with ballistic test results. The model will provide more flexibility in expanding tube design, leading to economic benefits in the overall expanding tube development procedure.

Demonstration of Compact Air Separation Technology for In-Flight Oxygen Collection Space Launchers

In-flight oxygen collection enables a significant reduction in takeoff gross weight of space launchers, thereby allowing airplane-like operations. This paper describes the development and tests of a compact rotary air separation device that constitutes a part of a functional oxygen collection plant. The experimental results are described in detail, demonstrating a consistent and long duration operation of the device yielding a high purity of liquid enriched air (96% oxygen) while causing only very small pressure drops on the gaseous phase. The tests indicate that the height of a so-called ideal tray for the current device is around 1 to 2 cm of core foam material, which is an improvement of one order of magnitude compared with previous designs. The experimental results are scaled up and applied to two different two stage to orbit launch vehicles. It is shown that a complete oxygen collection plant for a subsonic collection phase would weigh around 4 tons and occupy approximately , which represents less than 1% of the vehicle takeoff weight and a specific mass and volume, respectively, of and of liquid oxygen generated. For a supersonic collection, a much bigger collection plant is required due to a much higher oxygen production rate, and further improvements are required to reduce its system mass and volume despite a reduction in both specific mass and volume.

Software Fault Injection Testing of the Embedded Software of a Satellite Launch Vehicle

The software performing navigation, guidance, control, and mission-sequencing functionalities embedded in the flight computer system (FCS) of a satellite launch vehicle must be highly dependable. The presence of faults in the embedded flight software affects its dependability and may even jeopardize the entire mission, resulting in a huge loss to the space agency concerned. There are many techniques available to achieve high dependability and can be classified under fault avoidance, fault removal and fault tolerance.

Reduction of the actuator oscillations in the flying vehicle under a follower force

Flexible behaviors in new aerospace structures can lead to a degradation of their control and guidance system and undesired performance. The objectives of the current work are to analyze the vibration resulting from the propulsion force on a Single Stage to Orbit (SSTO) launch vehicle (LV). This is modeled as a follower force on a free-free Euler-Bernoulli beam consisting of two concentrated masses at the two free ends. Once the effects on the oscillation of the actuators are studied, a solution to reduce these oscillations will also be developed. To pursue this goal, the stability of the beam model is studied using Ritz method. It is determined that the transverse and rotary inertia of the concentrated masses cause a change in the critical follower force. A new dynamic model and an adaptive control system for an SSTO LV have been developed that allow the aerospace structure to run on its maximum bearable propulsion force with the optimum effects on the oscillation of its actuators. Simulation results show that such a control model provides an effective way to reduce the undesirable oscillations of the actuators.