Hybrid Rocket   :

            A hybrid rocket is a rocket with a rocket motor which uses propellants in two different states of matter - one solid and the other either gas or liquid. 

                Hybrid rockets exhibit advantages over both liquid rockets and solid rockets especially in terms of simplicity, safety, and cost. Because it is nearly impossible for the fuel and oxidizer to be mixed intimately (being different states of matter), hybrid rockets tend to fail more benignly than liquids or solids. Like liquid rockets and unlike solid rockets they can be shut down easily and are simply throttle-able. 

             The theoretical specific impulse,  Isp  performance of hybrids is generally higher than solids and roughly equivalent to hydrocarbon-based liquids.  Isp as high as 400s has been measured in a hybrid rocket using metalized fuels. Hybrid systems are slightly more complex than solids, but the significant hazards of manufacturing, shipping and handling solids offset the system simplicity advantages.

             In its simplest form a hybrid rocket consists of a pressure vessel (tank) containing the liquid propellant, the combustion chamber containing the solid propellant, and a valve isolating the two. When thrust is desired, a suitable ignition source is introduced in the combustion chamber and the valve is opened. The liquid propellant (or gas) flows into the combustion chamber where it is vaporized and then reacted with the solid propellant. Combustion occurs in a boundary layer diffusion flame adjacent to the surface of the solid propellant.

                  Generally the liquid propellant is the oxidizer and the solid propellant is the fuel because solid oxidizers are problematic and lower performing than liquid oxidizers. Furthermore, using a solid fuel such as Hydroxyl-Terminated Poly Butadiene (HTPB) or paraffin wax allows for the incorporation of high-energy fuel additives such as aluminium, lithium, or metal hydrides.

             Common oxidizers include gaseous or liquid oxygen or nitrous oxide. Common fuels include polymers such as polyethylene, cross-linked rubber such as HTPB or liquefying fuels such as paraffin wax.

Advantages compared with bipropellant Liquid Rockets : 

1. Mechanically simpler. 2. Denser fuels - fuels in the solid phase generally have higher density than those in the

liquid phase, reducing overall  system volume.

3. Metal additives - reactive metals such as aluminium, magnesium, lithium or beryllium can be easily included in the fuel grain increasing specific impulse, density specific impulse, or both.

Advantages compared with Solid Rockets :

1. Higher theoretical Isp  is possible.2. Less explosion hazard .3. More controllable .4. Relatively safe and non-toxic oxidizers such as liquid oxygen and nitrous oxide can be

used.5. Can be transported to site in a benign form and loaded with oxidizer remotely

immediately before launch, improving safety.

Disadvantages of Hybrid Rockets :

1. Oxidizer-to-fuel ratio shift ("O/F shift”).2. Low regression-rate .3. Compared with Liquid based propulsion, re-fuelling a partially or totally depleted hybrid

rocket would present significant challenges, as the solid propellant cannot simply be pumped into a fuel tank. 

# Cooling in Liquid Rocket :

            The heat created during combustion in a rocket engine is contained within the exhaust gases. Most of this heat is expelled along with the gas that contains it; however, heat is transferred to the thrust chamber walls in quantities sufficient to require attention.

                Regenerative cooling is the most widely used method of cooling a thrust chamber and is accomplished by flowing high-velocity coolant over the back side of the chamber hot gas wall to convectively cool the hot gas liner. The coolant with the heat input from cooling the liner is then discharged into the injector and utilized as a propellant.

              Earlier thrust chamber designs, had low chamber pressure, low heat flux and low coolant pressure requirements, which could be satisfied by a simplified "double wall chamber" design with regenerative and film cooling. 

            For subsequent rocket engine applications, however, chamber pressures were increased

and the cooling requirements became more difficult to satisfy. It became necessary to design new coolant configurations that were more efficient structurally and had improved heat transfer characteristics.

            This led to the design of "tubular wall" thrust chambers, by far the most widely used design approach for the vast majority of large rocket engine applications. These chamber designs have been successfully used and several other Air Force and NASA rocket engine applications. The primary advantage of the design is its light weight and the large experience base that has accrued. But as chamber pressures and hot gas wall heat fluxes have continued to increase (>100 atm), still more effective methods have been needed.

                One solution has been "channel wall" thrust chambers, so named because the hot gas wall cooling is accomplished by flowing coolant through rectangular channels, which are machined or formed into a hot gas liner fabricated from a high-conductivity material, such as copper or a copper alloy.. Heat transfer and structural characteristics are excellent.

            Basically there are three domains in a regenerative cooled rocket engine.

                           Gas Domain (Combusted Gases) - Convection and Radiation heat transfer                           Liquid Domain (Coolant) - Convection heat transfer                           Solid Domain (Thrust chamber wall) - Conduction heat transfer

            Heat transfer from the outer surface of thrust chamber to the environment can be neglected and the outer surface wall can be assumed as adiabatic.

                    In addition to the regenerative cooled designs mentioned above, other thrust chamber designs have been fabricated for rocket engines using dump cooling, film cooling, transpiration cooling, ablative liners and radiation cooling. Although regeneratively cooled combustion chambers have proven to be the best approach for cooling large liquid rocket engines, other methods of cooling have also been successfully used for cooling thrust chamber assemblies. 

               Dump cooling, which is similar to regenerative cooling because the coolant flows through small passages over the back side of the thrust chamber wall. The difference, however, is that after cooling the thrust chamber, the coolant is discharged overboard through openings at the aft end of the divergent nozzle. This method has limited application because of the performance loss resulting from dumping the coolant overboard. To date, dump cooling has not been used in an actual application.

               Film cooling provides protection from excessive heat by introducing a thin film of coolant or propellant through orifices around the injector periphery or through manifolded orifices in the chamber wall near the injector or chamber throat region. This method is typically used in high heat flux regions and in combination with regenerative cooling.

                   Transpiration cooling provides coolant (either gaseous or liquid propellant) through a porous chamber wall at a rate sufficient to maintain the chamber hot gas wall to the desired

temperature. The technique is really a special case of film cooling.

          With ablative cooling, combustion gas-side wall material is sacrificed by melting, vaporization and chemical changes to dissipate heat. As a result, relatively cool gases flow over the wall surface, thus lowering the boundary-layer temperature and assisting the cooling process.

            With radiation cooling, heat is radiated from the outer surface of the combustion chamber or nozzle extension wall. Radiation cooling is typically used for small thrust chambers with a high-temperature wall material (refractory) and in low-heat flux regions, such as a nozzle extension.

Nuclear Thermal Rocket  :

            In a nuclear thermal rocket a working fluid, usually liquid hydrogen, is heated to a high temperature in a nuclear reactor, and then expands through a rocket nozzle to create thrust.

Solid Core

            The most traditional type uses a conventional (albeit light-weight) nuclear reactor running at high temperatures to heat the working fluid that is moving through the reactor core. This is known as the solid-core design, and is the simplest design to construct.

           They are limited by the melting point of the materials used in the reactor cores. Since the efficiency of a rocket engine is related to the square root of the temperature of the working fluid, the solid core design needs to be constructed of materials that remain strong at as high a temperature as possible. 

           Nuclear reactions can create much higher temperatures than the temperatures the materials can withstand. Even more limiting is the cracking of fuel coatings due to the large temperature ranges (from 22 K up to 3000 K over the length of a 1.3m fuel rod), and the necessity of matching coefficients of expansion in all the components. Using hydrogen propellant, a solid-core design typically delivers specific impulses (Isp) on the order of 850 to 1000 seconds, about twice that of liquid hydrogen-oxygen designs such as the Space Shuttle Main Engine.

Liquid Core                      Dramatically greater improvements are theoretically possible by mixing the nuclear fuel into the working fluid and allowing the reaction to take place in the liquid mixture itself. This idea is the basis of the so-called liquid-core engine, which can operate above the melting point of the nuclear fuel--whatever the container wall can withstand while actively cooled by the hydrogen. The liquid-core design is expected deliver performance on the order of 1300 to 1500 seconds 

               These engines are currently considered to be very difficult to build. The reaction time of the nuclear fuel is much longer than the heating time of the working fluid and therefore requires a method to trap the fuel inside the engine while allowing the working fluid to easily exit through the nozzle. Most liquid-phase engines have focused on rotating the fuel/fluid mixture at very high speeds to press the fuel to the outside by centripetal force. 

Gas Core

            The final classification is the gas-core engine. This is a modification to the liquid-core design which uses rapid circulation of the fluid to create a toroidal pocket of gaseous uranium fuel in the middle of the reactor, surrounded by hydrogen. In this case the fuel does not touch the reactor wall at all, so temperatures could reach several tens of thousands of degrees, which would allow specific impulses of 3000 to 5000 seconds.

# Principles of Rocket Propulsion :            A rocket is a machine that develops thrust by the rapid expulsion of matter. The major components of a chemical rocket assembly are a rocket motor or engine, propellant consisting of fuel and an oxidizer, a frame to hold the components, control systems and a cargo such as a satellite. A rocket differs from other engines in that it carries its fuel and oxidizer internally, therefore it will burn in the vacuum of space as well as within the Earth's atmosphere. The cargo is commonly referred to as the payload. A rocket is called a launch vehicle when it is used to launch a satellite or other payload into space. A rocket becomes a missile when the payload is a warhead and it is used as a weapon. At present, rockets are the only means capable of achieving the altitude and velocity necessary to put a payload into orbit.

Terms to Describe Rocket Power    

            There are a number of terms used to describe the power generated by a rocket.

                      Thrust is the force generated, measured in pounds or kilograms. Thrust generated by the first stage must be greater than the weight of the complete launch vehicle while standing on the launch pad in order to get it moving. Once moving upward, thrust must continue to be generated to accelerate

the launch vehicle against the force of the Earth's gravity. To place a satellite into orbit around the Earth, thrust must continue until the minimum altitude and orbital velocity have been attained or the launch vehicle will fall back to the Earth. Minimum altitude is rarely desirable, therefore thrust must continue to be generated to gain additional orbital altitude.

              The impulse, sometimes called total impulse, is the product of thrust and the effective firing duration. A shoulder fired rocket such as the LAW has an average thrust of 600 lbs and a firing duration of 0.2 seconds for an impulse of 120 lb-sec. The Saturn V rocket, used during the Apollo program, not only generated much more thrust but also for a much longer time. It had an impulse of 1.15 billion lb¬sec.

                      The efficiency of a rocket engine is measured by its specific impulse (Isp). Specific impulse is defined as the thrust divided by the mass of propellant consumed per second. The result is expressed in seconds. The specific impulse can be thought of as the number of seconds that one pound of propellant will produce one pound of thrust. If thrust is expressed in pounds, a specific impulse of 300 seconds is considered good. Higher values are better.

                A rocket's mass ratio is defined as the total mass at lift-off divided by the mass remaining after all the propellant has been consumed. A high mass ratio means that more propellant is pushing less launch vehicle and payload mass, resulting in higher velocity. A high mass ratio is necessary to achieve the high velocities needed to put a payload into orbit.

# Ramjet Inlet Operation :

               There are three distinct conditions under which a ramjet engine diffuser can operate, depending on the heat released in the combustor. 

Critical

               When the heat released in the combustor is just enough that the back pressure at the exit section of the subsonic diffuser causes the normal shock to be positioned at the inlet throats, the operation is said to be critical; this is the design condition. 

Subcritical  

               If the heat release in the combustor is increased, the static pressure at the exit of the subsonic diffuser is greater than can be achieved under the design condition. The normal shock wave moves upstream, is expelled from the diffuser, and continues to move toward the vertex of the supersonic diffuser. Behind the normal shock wave, the flow is subsonic. Since the shock wave is detached from the inlet, the incoming air spills over the cowl of the diffuser.

increasing vehicle drag and possibly leading to instability. 

Supercritical

               When the heat released in the combustor is decreased, the back pressure at the outlet section of the diffusion system becomes too small to maintain the normal shock at the inlet. The excess pressure associated with the internal flow must therefore be dissipated inside the diffusion system by a strong shock wave forming in the diverging portion of the diffuser. In other words the normal shock moves into the inlet. 

            These three operating conditions can be related conveniently by means of pressure recovery and  mass flow rate of air.

Thrust Vector Control :

            Thrust vector control is effective only while the propulsion system is creating thrust. At other stages of flight, separate mechanisms are required for attitude and flight path control.

           Nominally, the line of action of the thrust vector of a rocket nozzle passes through the vehicle's center of mass, generating zero net moment about the mass center. It is possible to generate pitch and yaw moments by deflecting the main rocket thrust vector so that it does not pass through the mass center. Because the line of action is generally oriented nearly parallel to

the roll axis, roll control usually requires the use of two or more separately hinged nozzles or a separate system altogether, such as fins, or vanes in the exhaust plume of the rocket engine, deflecting the main thrust.

                Thrust vectoring for many liquid rockets is achieved by gimballing the rocket engine. This often involves moving the entire combustion chamber and outer engine bell as on the Titan II's twin first stage motors, or even the entire engine assembly including the related fuel and oxidizer pumps. Such a system was used on the Saturn V and the Space Shuttle.

      Another method of thrust vectoring used on early solid propellant ballistic missiles was liquid injection, in which the rocket nozzle is fixed, but a fluid is introducedinto the exhaust flow from injectors mounted around the aft end of the missile. If the liquid is injected on only one side of the missile, it modifies that side of the exhaust plume, resulting in different thrust on that side and an asymmetric net force on the missile. This was the control system used on the Minuteman II and the early SLBMs of the United States Navy.

               A later method developed for solid propellant ballistic missiles achieves thrust vectoring by deflecting the rocket nozzle using electric servo mechanisms or hydraulic cylinders. The nozzle is attached to the missile via a ball joint with a hole in the center, or a flexible seal made of a thermally resistant material, the latter generally requiring more torque and a higher power actuation system. The Trident C4 and D5 systems are controlled via hydraulically actuated nozzle.

#  Scramjet  :

         A scramjet (supersonic combustion ramjet) is a variant of a ramjet air-breathing jet engine in which combustion takes place in supersonic airflow. 

              As in ramjets, a scramjet relies on high vehicle speed to forcefully compress and decelerate the incoming air before combustion (hence ramjet), but whereas a ramjet decelerates the air to subsonic velocities before combustion, airflow in a scramjet is supersonic throughout the entire engine. 

        This allows the scramjet to operate efficiently at extremely high speeds: theoretical projections place the top speed of a scramjet between Mach 12 and Mach 24.

Basic Components :

          The scramjet is composed of three basic components.

          A converging inlet, where incoming air is compressed and decelerated.           A combustor, where gaseous fuel is burned with atmospheric oxygen to produce heat.          A diverging nozzle, where the heated air is accelerated to produce thrust. 

       Unlike a typical jet engine, such as a turbojet or turbofan engine, a scramjet does not use rotating, fan-like components to compress the air; rather, the achievable speed of the aircraft moving through the atmosphere causes the air to compress within the inlet. 

          As such, no moving parts are needed in a scramjet. In comparison, typical turbojet engines require inlet fans, multiple stages of rotating compressor fans, and multiple rotating turbine stages, all of which add weight, complexity, and a greater number of failure points to the engine.

          Due to the nature of their design, scramjet operation is limited to near-hypersonic velocities. 

      As they lack mechanical compressors, scramjets require the high kinetic energy of a hypersonic flow to compress the incoming air to operational conditions. 

          Thus, a scramjet-powered vehicle must be accelerated to the required velocity by some other means of propulsion, such as turbojet, railgun, or rocket engines.  

Advantages

1.   Does not have to carry oxygen.2.   No rotating parts makes it easier to manufacture.3.   Has a higher specific. 4.   Higher speed could mean cheaper access to outer space in the future.

Disadvantages  

1.    A scramjet cannot produce efficient thrust unless boosted to high speed, around Mach 5,         although depending on the design it could act as a ramjet at low speeds. 

2.    Testing scramjet designs use extremely expensive hypersonic test chambers or expensive      

        launch vehicles, both of which lead to high instrumentation costs. 3.    Lack of stealth.

# Solar Sail :

          A solar sail is a spacecraft propelled by sunlight. Whereas a conventional rocket is propelled by the thrust produced by its internal engine burn, a solar sail is pushed forward simply by light from the Sun. This is possible because light is made up of packets of energy known as “photons” that act like atomic particles, but with more energy. 

           When a beam of light is pointed at a bright mirror-like surface, its photons reflect right back, just like a ball bouncing off a wall. In the process the photons transmit their momentum to the surface twice – once by the initial impact, and again by reflecting back from it. Ever so slightly, propelled by a steady stream of reflecting photons, the bright surface is pushed forward.

           A solar sail is made up of just such a reflective surface, or several surfaces, depending on the sail’s design. When the bright sails face the Sun directly, they are subjected to a steady barrage of photons that reflect off the shiny surfaces and impel the spacecraft forward, away from the Sun. By changing the angle of the sail relative the Sun it is possible to affect the direction in which the sail is propelled – just as a sailboat changes the angle of its sails to affect its course. It is even possible to direct the spacecraft towards the Sun, rather than away from it, by using the photon’s pressure on the sails to slow down the spacecraft’s speed and bring its orbit closer to the Sun.

           In order for sunlight to provide sufficient pressure to propel a spacecraft forward, a solar sail must capture as much Sunlight as possible. This means that the surface of the sail must be big – very big. Even with such a gigantic surface, a solar sail spacecraft will accelerate very slowly when compared to a conventional rocket. Under optimal conditions, a solar sail on an interplanetary mission would gain only 1 millimeter per second in speed every second it is pushed along by Solar radiation. The Mars Exploration Rovers, by comparison, accelerated by as much as 59 meters per second every second during their launch by conventional Delta II rockets. This acceleration is 59,000 times greater than that of a solar sail.

            But the incomparable advantage of a solar sail is that it accelerates CONSTANTLY. A rocket only burns for a few minutes, before releasing its payload and letting it cruise at a constant speed the rest of the way. A solar sail, in contrast, keeps on accelerating, and can ultimately

reach speeds much greater than those of a rocket-launched craft. 

Operation

            Sailing operations are simplest in interplanetary orbits, where attitude changes are done at low rates. For outward bound trajectories, the sail force vector is oriented forward of the Sun line, which increases orbital energy and angular momentum, resulting in the craft moving farther from the Sun. For inward trajectories, the sail force vector is oriented behind the Sun line, which decreases orbital energy and angular momentum, resulting in the craft moving in toward the Sun. It is worth noting that only the Sun's gravity pulls the craft toward the Sun.

Aerospike Nozlle

Nozzle

Pressure and Pump Fed Rocket

Solid and Liquid Rocket

Solid Grain Configuration

Solid Propellant Rocket