Rocket Propulsion

Carbon-based caged and heterocyclic compounds tend to have strained molecular structures leading to high heats of formation and energetic behavior. In the current paper, molecular modelling calculations for 10 caged compounds of this type along with 2 strained aliphatic compounds and 4 simple cyclic chains are presented in view of their possible use as oxidizers in propulsion applications. Density functional theory (B3LYP) was employed for the geometry optimization of the proposed molecular structure using the 6-311++G(d,p) basis set. Heats of formation of the compounds were calculated using the molecular modeling results and their specific impulses were computed using the NASA CEA [1] software package to evaluate their potentials as propellant oxidizers.

Momentarily Displacement Force Thrust the momentarily displacement force thrust (to be referred to as MDF thrust) is a closed system that doesn't use any rocket propulsion nor propels any matter, molecules or ions..etc to produce thrust. The system facilitates the theory/phenomenon of the "momentarily displacement force" (not yet proven or put into experiment) resulting from releasing less dense matter (floats, air bubbles...etc) from under the surface of a higher density fluid. A. Momentarily Displacement Force (hypothesis) any matter ingested in fluid container will cause the fluid level to rise, increasing the pressure on the container's floor, when that pressure is multiplied to the floor area, that would be the forces towards the acceleration (gravity).

A compact aneutronic fusion engine will enable more challenging exploration missions in the solar system. This engine uses a deuterium-helium-3 reaction to produce fusion energy by employing a novel field-reversed magnetic field configuration (FRC). The FRC has a simple linear solenoidal coil configuration yet generates higher plasma pressures for a given magnetic field than other designs. Waste heat generated from bremsstrahlung and synchrotron radiation is recycled to maintain the fusion temperature. The charged reaction products, augmented by additional propellant, are exhausted through a magnetic nozzle. As an example, we present a mission to deploy the James Webb Space Telescope from LEO to an L2 halo orbit using a one MW compact aneutronic fusion rocket engine. The engine produces 20 N of thrust with an exhaust velocity of 55 km/s and has a specific power of 0.70 kW/kg.

NASA's JUNO mission will arrive at Jupiter in July 2016, after nearly five years in space. Since operational costs tend to rise with mission time, minimizing such times becomes a top priority. We present the conceptual design for a 10MW aneutronic fusion engine with high exhaust velocities that would reduce transit time for a Jupiter mission to eighteen months and enable more challenging exploration missions in the solar system and beyond. __________________________________________________

The object moving at high speeds encounter forces which tend to decelerate the objects. This resistance in the medium is termed as drag which is one of the major concerns while designing high speed aircrafts. Another key factor which influences the design is the air drag. The main challenges faced by aerospace industries is to design the shape of nose cone of flying object that travels at high speeds with optimum values of air drag. This study deals with computational analysis on elliptical nose cone profiles of a commercial aircraft. The effect of nose shape on the drag is studied at subsonic flow with Mach 0.3.The scope of this paper is to develop some prototype profiles with outstanding aerodynamic qualities and low cost for use in construction projects for aircrafts increasing their efficiency and effect on target. The motivation for such a work is caused by lack of data on aerodynamics for profile of elliptical nose cone and especially improved aerodynamic qualities that can be used in designing aircrafts. Flow phenomena observed in numerical simulations for different AOA for elliptical nose cone profile are highlighted, critical design aspects and performance characteristics of selected nose cone are presented.

This work is devoted to the study and the understanding of the performances of solid propulsion. This study focuses on two essential points: the energy performances study, which is based on the combustion study of the most widely used solid propellant (AP/HTPB/Al) in order to determine the thermochemical parameters using different solid propellant geometries which we compare their effects on the performance of solid propellants. The study of propulsion
performances is the second point and it is based on the study of the behavior of gases in the convergent divergent nozzle which was realized by a simulation on Ansys Fluent 19.1.
Key words: solid propellants, solid propulsion, combustion, nozzle

The industrial and scientific communities are devoting major research efforts to identify and assess critical technologies for new advanced propulsive concepts: combustion at high pressure, as well as the replacement of hydrogen with a hydrocarbon, has been assumed as a key ...

Представлены результаты численного моделирования рабочего процесса в ракетном двигателе нового типа на криогенном твердом топливе (КТТ), которое содержит горючее и окислитель при температуре ниже их температуры плавления. Математическая модель рабочего процесса базируется на полной системе уравнений Навье – Стокса для газа (с учетом k–ε-модели тубулентности) и уравнениях тепло- и массопереноса для твердого топлива, записанных в нестационарной осесимметричной форме. Расчеты выполнены для КТТ кислород + водород и кислород + метан. Рассмотрены дисковая и матричная схемы расположения топливных элементов канального заряда и проанализировано влияние их размеров на протекание рабочего процесса. Сделан вывод о невозможности реализации устойчивого горения топлива кислород + водород в данных условиях, однако показано, что использование топлива кислород + метан дает возможность реализации устойчивого рабочего процесса в камере сгорания и управления скоростью преобразования компонентов топлива путем задания определенной внутренней структуры КТТ. Рассмотрены также особенности рабочего процесса в гибридном ракетном двигателе принципиально нового класса на криогенном топливе, один из компонентов которого находится в жидком, а другой — в твердом агрегатном состоянии.

Space missions being incredibly expensive, the industry is always looking for ways of reducing the costs. A huge percentage of the cost of any space mission is the launch. Therefore, the recent technology of reusing launch vehicles has been revolutionary. This report focuses on the control surfaces of such reusable launch vehicles with lattice fins in particular. After conducting wind tunnel testing for several fin designs, a conclusion as to the best available design will be made.

Linear aerospike nozzles are studied here, as a possible device to improve launcher engine performance. The main property of these nozzles is the possibility of good integration with vehicle. To improve the knowledge of flow-field and spike pressure distribution a 3-D aerospike nozzle is studied numerically and experimentally at atmospheric conditions and by varying chamber conditions. The spike plate pressure variation is measured by attaching pressure sensors at four different points of the plate and is compared with CFD results.

In the first decades of space exploration several concepts appeared for utilizing nuclear energy in space. However, only a few of them were developed into actual instruments. Radioisotope powered batteries and heating units have been regularly used in automatized space probes since then. Small nuclear reactors were also launched into space in the Cold War era, but the hazards finally proved to be too great compared to the benefits. Nuclear propulsion systems were also developed in the USA (NERVA) and in the USSR (RD 0410) but never actually used in a launched spacecraft. The concept of a radioisotope powered thermal rocket engine also existed, but the initial experiments revealed several problems, so the concept was set aside. Nuclear energy in space Cascade decaying isotopes Outline of a possible implementation of the THOR engine (Thorium-228 isotope powered thermal rocket engine). The body of the engine could be made of tantalum-hafnium carbide ceramic, which has a melting point at 4260 K-this material allows high working temperatures (3300-3600 K) while effectively absorbing the radiation of the source. The hydrogen propellant leaves the nozzle at a speed of 10 km/sec at this temperature. Adding a small amount of hydrocarbons to the propellant material prevents the chemical wear of the engine parts by the hot hydrogen gas flowing through them.

Retrogradation rate and extent were determined by differential scanning calorimetry and rheology measurements on 40% dry matter gels made from pea, modified waxy maize, rice, waxy rice, wheat, manioc, potato and microwaved irradiated potato starches. As a result, each starch retrogrades differently, depending also on the measurement technique. Nevertheless, potato and pea starches seem to be the most sensitive and the waxy and modified waxy types are the least sensitive to retrogradation. ~ ~~ Retrogradationskinetik von acht Starken. Retrogradationsgeschwindigkeit und -ausmaR wurden mittels Differential-Rastercalorimetrie und rheologischen Messungen an Gelen mit 40% TS aus Erbsen-, modifizierter Wachsmais-, Reis-, Wachsreis-, Weizen-, Maniok-, Kartoffel-und mikrowellenbestrahlter Kartoffelstarke bestimmt. Es zeigte sich, daR alle Starken unterschiedlich retrogradieren, auch in Abhangigkeit von der MeBtechnik. Trotzdem scheinen Kartoffel-und Erbsenstarken die empfindlichsten und die wachsigen und modifizierten wachsigen Arten die am wenigsten empfindlichen Starken gegeniiber der Retrogradation zu sein. ~ starctdstarke 42 (1990) Nr. 3. S . 99-101 0 VCH Verlagsgesellschaft mbH, D-6940 Welnheim. 1990 0038-9056/90/0303-09$02.50/0

در طراحی سامانههای پرتاب پیشرانه جامد)سوخت جامد( پارامترهای عملکردی زیادی در روند کار دخیل میباشد. از
مهمترین مسائل طراحی موتورهای راکتی و موشکی پیشرانه جامد، بررسی مشخصه و نوع ماده پیشران است. در الگوریتم
طراحی موتورهای پیشران جامد، اولین قدم شناخت و انتخاب نوع ماده پیشران و به کارگیری مولفهها و پارامترهای اصلی
طراحی است. شناخت جامع این مشخصهها به لحاظ تئوری و تجربی جلوه کاملی به کل سامانه میبخشد. خواص فیزیکی
و مکانیکی پیشرانهها، روابط و مشخصههای موادی پیشرانهها به همراه نحوه تحلیل در مشخصههای گرمایی و احتراق ماده
پرانرژی رویکرد این مقاله خواهد بود. در انتها بیشتر خواص بیان شده خود را در پارامتر نرخ سوزش نشان میدهند که به
نحوه اندازهگیری آن با روشهای تجربی و تئوری پرداخته شده است. با تمام این اوصاف هدف مدنظر و نوع ماده پیشرانه
در تمام مراحل طراحی و تحلیل موتورهای پیشران جامد دخیل خواهد بود و یک پارچهسازی تئوری با موارد تجربی موجب
تعیین و پیشبینی صحیح مولفههای عملیاتی موتور پیشران جامد خواهد شد.

This is a particular section of the author's propulsion lab (group) project of which the author contributed the most in the making of the report. In general, the calculations that have been made in the consideration of the rocket aerodynamics can be differentiated into two different approaches – static & dynamics. The static takes into account the first moments, centroids & center of gravity. Whereas, the dynamics consists of peak height, the velocity of rocket, the time lapse, drag equation, the nozzle thrust & dynamic pressure. Nonetheless, most of the equations have been considered with few assumptions made. It is for certain that any possible discrepancy is inevitable in this particular section in comparison of the actual results obtained. Lastly, the results obtained are as follows: centre of gravity = 19 cm upward with respect to the nozzle-end; centre of pressure = 5 cm upward with respect to the nozzle-end; peak height = 33.69 cm; instantaneous velocity = 25.71 m/s; time of free-falling = 2.62 s; time of upward motion = 2.38 s; drag = 1.21 N; instantaneous nozzle thrust = 608.4 N; and dynamic pressure = 397.9 Pa.

One of the principal limitations of long duration human spaceflight beyond cis-lunar orbit is the lack of refuelling capabilities on distant planets resulting in the reliance on conventional non-cryogenic, propellants produced on Earth. If one could develop a reliable propulsion system operating on propellants derived entirely of ingredients found on nearby planetary bodies, then not only could mission duration be extended, larger amounts of payload could be ferried to and from the destination and eventually the cost of transporting propellant ingredients from Earth could be reduced, if not eliminated. Metal powder, in particular, aluminium burns energetically with oxygen and affords an impressive exothermic reaction that can be utilised for propulsive force. The oxidiser component of an aluminium powder based propellant can be in the form of pure oxygen, air, steam or a compound such as ammonium perchlorate (AP) making it ideal for terrestrial applications. So much so, that aluminium powder as a fuel has been investigated intermittently for over sixty years for military applications such as to propel torpedoes using an oxidiser in the form of steam produced by drawing in seawater, ram-air as an oxidiser for air-launched ram jet missiles and AP in powder form which when combined and ignited with aluminium powder results in combustion which can deliver a propulsive force suitable for long range surface-to-surface missiles. In terms of space applications , the logic of employing oxygen/aluminium metal powder as an in-situ resource utilisation (ISRU) propulsion concept is due to the fact that the propellant combination are constituent chemicals that can be found on numerous planetary bodies and thus existing research could be adapted and enhanced for possible employment in space based systems. Therefore the technological readiness level is not as complex as if one were to develop an entirely new propulsion system. Living off the land to realise extended duration residency on extra-terrestrial bodies of interest is not exactly a revelation. Metal powder propulsion systems have been addressed intermittently since the Second World War, initially in the field of underwater propulsion where research in the application of propelling torpedoes continues until this day. During the post war era, researchers attempted to utilise metal powders as a fuel for ram jet applications in missiles. The 1960's and 1970's saw additional interest in the use of 'pure powder' propellants, i.e. fluidised metal fuel and oxidiser, both in solid particulate form. Again the application was for employment in space-constrained missiles where the idea was to maximise the performance of high energy density powder propellants in order to enhance the missile's flight duration. Metal powder as possible fuel was investigated for in-situ resource utilisation propulsion systems post-1980's where the emphasis was on the use of gaseous oxygen or liquid oxygen combined with aluminium metal powder for use as a " lunar soil propellant " or carbon dioxide and magnesium metal powder as a " Martian propellant ". Albeit aluminium metal powder propellants are lower in performance than cryogenic and Earth storable propellants, the former does have an advantage inasmuch that the propulsion system is generic, i.e. it can be powered with chemicals mined and processed on Earth, the Moon and Mars. Thus, due to the potential refuelling capability, the lower performing aluminium metal powder propellant would effectively possess a much higher change in velocity (∆V) for multiple missions than the cryogenic or Earth storable propellant which is only suitable for one planet or one mission scenario, respectively. This paper reviews existing metal powder propulsion systems, while emphasising the option of oxygen/aluminium metal powder as a propellant, in order to highlight the potential of this near-term concept.

Of all the planets of our Solar system, Mars is by far the most similar one to Earth, despite it being an inhabitable, radiation-rich land with a toxic atmosphere. Its position with respect to the Sun, its rocky environments and its mass - all of which may have, a long time ago, helped it host liquid water and some form of ancient life - make the Red Planet the perfect candidate for the future of manned missions.
As a matter of fact, having a permanent ground station on in-between planets and satellites, like our Moon and Mars, may be a giant leap towards space exploration. This may lead to a better understanding of our Universe. However, in order to reach that stage of human civilization, many small steps are yet to be taken. Firstly, we should prioritize the protection and well being of astronauts during months long travels in micro-gravity. Next, we should focus on the protection and well being of astronauts during their permanence on the celestial body. Finally, we should ensure astronauts have a safe return trip, via secure fuel provisioning.
Nevertheless, these goals are not easy to achieve, especially due to the fact that not every resource needed for survival can be brought on board, given the limited capacity of the spaceship and the extremely high cost to weight ratio of a single launch.
In fact, it would be much cheaper to manipulate some of the resources already present on Mars (or the Moon) so as to create all that is needed from scratch.


This paper will discuss the methodologies that can be used to manipulate Mars’ natural resources, a practice known as In-Situ Resource Utilization (ISRU). In the coming years, ISRU will be the main focus of many aerospace companies, as it constitutes the next step towards human colonization of Mars and the Moon.
The purpose of this paper is to suggest an indicative guideline for the exploitation of most martian resources. At the start, this will be carried out by using automated equipment. Then, it will continue via human-run hardware installed in-situ. This will be necessary in order to build a livable habitat and fabricate fuel for the return trip.
Given the low technology readiness level (TRL) of numerous technological applications that this paper will explore, some of the solutions to the challenges that ISRU puts before engineers will be discussed in qualitative rather than quantitative terms.
A strong emphasis will be put in the design of the power system of a hypothetical martian habitat using exclusively photovoltaic panels, as this is the most representative topic of my course of studies.

A complete theoretical and mathematical description of how electric fields change due to relative motion that includes an in-depth analysis of relativistic electric fields and the methods to effect motion.

This new description is based on James Maxwell's original bi-quaternion that includes descriptions and examples of how to use these electric field differences to create an all new propellant-less spacecraft propulsion device.

Additive Manufacturing (AM) of metals is a processing technology that has significantly matured over the last decade. For liquid propellant rocket engines, the advantages of AM for replacing conventional manufacturing of complicated and expensive metallic components and assemblies are very attractive. AM can significantly reduce hardware cost, shorten fabrication schedules, increase reliability by reducing the number of joints, and improve hardware performance by allowing fabrication of designs not feasible by conventional means. The NASA Marshall Space Flight Center (MSFC) has been involved with various forms of metallic additive manufacturing for use in liquid rocket engine component design, development, and testing since 2010. The AM technique most often used at the NASA MSFC has been powder-bed fusion or selective laser melting (SLM), although other techniques including laser directed energy deposition (DED), arc-based deposition, and laser-wire cladding techniques have also been used to develop several components. The purpose of this paper is to discuss the various internal programs at the NASA MSFC using AM to develop combustion devices hardware. To date at the NASA MSFC, combustion devices component hardware ranging in size from 100 lbf to 35,000 lbf have been designed and manufactured using SLM and deposition-based AM processes, and many of these pieces have been hot-fire tested. Combustion devices component hardware have included thrust chamber injectors, injector components such as faceplates, regeneratively-cooled combustion chambers, regeneratively-cooled nozzles, gas generator and preburner hardware, and augmented spark igniters. Ongoing and future developments for combustion devices have also included design of components sized for boost-class engines. Several design and hot-fire test iterations have been completed on these subscale and larger scale components, and a summary of these results will be presented as well.

Additive manufacturing is increasing in prevalence and
importance across many industries. This tutorial will provide a
thorough introduction to the technologies involved, the typical
manufacturing process, and common design considerations using
this new technology.

Cleaning has become a basic need for all human beings and it is an unavoidable daily routine process. A road cleaning machine has a wide range of applications. To overcome the rising issues of cleanliness in India the road cleaner can play a vital role. The conventional road cleaning machine is most widely used in railway stations, airports, hospitals, Bus stands, etc. It is not user friendly as well as Eco-friendly. In the summertime, there is a power crisis and most of the roads cleaning machines are not used effectively due to this problem particularly.
In our project, we are using easily available materials at low cost. The machine is an assembly of various rigid components. In this report, our motive is to present a detailed qualitative study of cleaning system used in our machine, the main focus being cleanliness with minimum utilization of resources available with us.

Shortly before a test launch of the first Russian R-16 intercontinental rocket at the Baikonur Cosmodrome on October 24, 1960; a terrible accident – the deadliest in the history of the space industry – claimed an estimated 126 lives, although some sources contend the actual number of deaths may be as high as 500. This paper describes that while on the launch stand undergoing last-minute repairs and preparations, a major malfunction in an electrical circuit caused a premature ignition of the second stage engines – detonating the fully loaded first stage fuel tanks. Also discussed are historical accounts of the decisions by Marshal Mitrofan Nedelin to shortcut safety procedures and continue with the test launch to meet arbitrary schedule demands from Moscow, resulting in a significant number of deaths, including that of Marshal Nedelin, and a total loss of the rocket and test stand. The paper provides historical context to explain the unrelenting schedule pressure. In addition, the paper describes the context of the Soviet space program at the time of the accident and provides a chronological outline of the events surrounding the failed test launch. Translating several sources from their original Russian, we present much of the story of the Nedelin Disaster through first-hand accounts and memoirs of eyewitnesses to the catastrophe, their survivors and other Soviet rocket and military officials. We present the mistakes, poor decisions and incredible lack of site planning that directly contributed to both the cause and the severity of the disaster. Throughout the paper, authoritative historical sources, many of them translated from Russian into English for what we believe is the first time, describe the events leading up to the accident, the devastation of the event, and the decades of duplicitous deception perpetuated by the Soviet government to hide the true story from the families of the victims, the Soviet people and the rest of the world. The paper's conclusion presents several lessons learned from this incident.

Global search is on for a solid fuel for the area weapon
Multi Barrel Rocket Launcher (MBRL). Solid fuel is the first
choice due to ease of design, manufacturing, handling, packaging,
transportability and deployability. It also has the capability of
prolonged storage without noticeable degradation of its quality. It
is also quite safe and does not pose safety hazards. Also with solid
propellant end use by the troops is comparatively simple as only
trigger operation is needed. Tactically it suits the requirement of
very fast response time and generates high thrust for small
duration. Solid fuel, better known as solid propellant is classified
as single base and double base propellant. Single base propellant
contains nitrocellulose (NC) where as double base contains
homogeneous mixture of nitrocellulose and nitroglycerine (NG)
with small amount of additives. There is a third category of solid
propellant known as composite propellant. Double base
propellant used to be preferred in tactical missiles and rockets.
However, with the advent of the third category, it is now most
preferred. In this category additives are added to meet the
requirement of manufacturing, operational requirements and to
improve physical properties. It contains heterogeneous mixture
of solid fuel and oxidizer held together in a matrix like binder.
Polybutadiene is the most effective binder which improves the
physical properties. In this mixture when solid ammonium
perchlorate (AP) and aluminum (Al) are added the density of the
mixture is increased resulting in improvement of specific impulse.
Presently composite propellant is cast from a mixture of AP
crystals, Al powder and liquid HTBP (Hydroxyl terminated
Polybutadiene). It is called CMDB (Composite Modified Double
Base) propellant. Pressure time curve for burning of CMDB
propellant obtained during static trial in Ballistic Evaluation
Motor establishes that propellant grain with star shape internal
and cylindrical outer cross section maintains constant pressure
over the entire burning time. This is because such cross section
provides constant (neutral) burning and leaves negligible sliver at
the end. This paper brings out recent advances in solid fuel and
analyses suitability of CMDB propellant to fulfill the requirement
for use in MBRL. From the analysis of Pressure-Time plot of the
subject propellant it is found to be most suitable as solid rocket
propellant for MBRL.

The HyShot 2 flight in 2002 pioneered a low cost method of hypersonic testing using sounding rockets. In a further development of this technology, Flight 7 of the HIFiRE Program will deploy a free-flying scramjet powered payload that is planned to enter the atmosphere at a high flight path angle at approximately Mach 8. The HIFiRE 7 payload consists of two back-to-back three-dimensional scramjet engines set within a trailing flare for aerodynamic stability. The scientific goals of the flight are to perform a direct measurement of the thrust generated by the scramjet flowpaths, and to compare the measured thrust with estimates based on ground testing. This paper describes the motivation, launch procedure, flowpath development and ground testing for the flight which is scheduled for March 2011.

Link to my video about the derivation of the rocket equation.
References:
* SPIEGEL, Murray R. Schaum's outline of theory and problems of theoretical mechanics: with an introduction to Lagrange's equations and Hamiltonian theory. McGraw-Hill Companies, 1967.
*HALLIDAY, David; RESNICK, Robert; WALKER, Jearl. Fundamentals of physics. John Wiley & Sons, 2013.

The article presented here consists of design and simulation process carried out for designing a two stage sounding rocket. The design and simulation of the sounding rocket was done in open source software, called OpenRocket. The final design was named as PRASOM sounding rocket, which are first three initials of team members. The PRASOM sounding rocket is a reusable type. From the simulation ran for PRASOM sounding rocket, its apogee is found to be around 3400 m. A 5 Kg payload weight was considered before starting the design process. This kind of payload can do the meteorological missions as well as behavioural biology (Collette and Kempf, 2007) missions. The total time of flight found from simulation is 215 s. The various specifications for PRASOM sounding rocket was made to freeze by review/analysis of previous Cal Poly Space System (CPSS) high-altitude rockets, Rohini RH200/125 sounding rocket and Atea-1 sounding rocket.

ACCREDITED by AICTE-NBA Maisammaguda, Dhulapally post, Secunderabad-500014 DECLARATION I hereby declare that the mini project entitled CFD Analysis of a Rocket Nozzle with one Inlet at Mach 0.6 submitted to Malla Reddy College of Engineering and Technology, affiliated to Jawaharlal Nehru Technological University Hyderabad (JNTUH) for the award of the degree of Bachelor of Technology in Aeronautical Engineering is a result of original research work done by us.

Design and Construction of a Fast Generational Starship
“On a recent visit to NASA Marshall Space Flight Center, NASA Glenn Research Center and the Tennessee Valley Interstellar Workshop, I challenged people to refute my controversial (and deliberately provocative) claim that Daedalus is the only starship design in history. NASA appeared to agree with me. This is an astonishing revelation and I find it intriguing that people are prepared to pronounce interstellar flight impossible when we have only attempted one such design in history. More feasibility studies are clearly required before we can have a clear picture of what the impossibilities or otherwise are.” Kelvin F. Long
The problem with this statement is that it is not actually true. Daedalus was designed to use Pulse Fusion, and that hasn’t panned out despite decades of research. You can’t really have a starship design without a workable propulsion system. Additionally, Daedalus was not a “starship” in the normal sense of the word but a giant unmanned space probe. That said, he is correct to point out that we have no idea what the real possibilities are because we haven’t really considered what those possibilities might be. This paper will explore the design of a Generational Starship using Fission Fragment (FF) propulsion with a Nuclear Pulse (NP) Booster. A Generational Starship defined as being a Starship that can travel to another star system within a human lifetime. The proposed Starship will carry a crew of four on a mission to Alpha Centauri (4.367ly) that will take forty years. Keeping the crew small substantially reduces the size of the Starship compared to past concepts for very large multi-generational Starships. Below is a conceptual view of what such a starship might look like.


Figure 1Concept Starship

Ground testing and flight testing- types of tests, test facilities and safeguards, monitoring and control of toxic materials, instrumentation and data management. Ground testing, flight testing, trajectory monitoring, post accident procedures. Description of a typical space launch vehicle launches procedure.

Cryogenics originated from two Greek words 'kayos' which means 'cold or freezing' and 'genes' which means 'born or produced'. Cryogenic is the study of very low temperatures or the production of the same. This paper is all about the rocket engine involving the use of cryogenic technology at a cryogenic temperature (123K). This basically uses the liquid oxygen and liquid hydrogen as an oxidizer and fuel, which are very clean and non-pollutant fuels compared to other hydrocarbon fuels like: petrol, diesel, gasoline, LPG, NCG etc., sometimes, liquid nitrogen is also used as an fuel. The efficiency of the rocket engine is more than the jet engine. As per the Newton's third law of mechanics, the thrust produced in the rocket engine is outwards whereas that produced in jet engine is inwards.

The first known reference to the use of rocket propulsion dates back to 1232 AD, when fire arrows were used by the Chinese for fighting off the Mongols. These ancient rockets used gunpowder as fuel. Since then, tremendous technological advancement has led rocket propulsion to ...