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Introduction to Astronautics Sissejuhatus kosmonautikasse

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1 Introduction to Astronautics Sissejuhatus kosmonautikasse
Tallinn University of Technology Introduction to Astronautics Sissejuhatus kosmonautikasse Vladislav Pustõnski 2009

2 Principle systems of rockets
A launch vehicle (and, in general, a rocket vehicle) is a very complex device containing a lot of systems, and normal flight of the rocket depends on their concerted work. We have already studied engines, navigation and guidance systems of rockets, now let us look in detail other important systems. Tanks Tanks are the largest assemblies of each launch vehicle, most of spacecrafts also have propellant tanks. We already know the principle requirements which should be fulfilled: the tanks should be as light as possible so that the mass of the structure remain small. In the case of solid rocket motors or pressure-fed liquid propellant engines the tanks should withstand high pressure (higher than that is present in the thrust chamber). There are two principle types of tanks: with a frame structure and a monocoque (without framework). The latter is also known as balloon tank, since it keeps its shape and stability due to internal pressure. This construction may be very light, it was used, for example, on the Atlas family rockets, except for its last member, the Atlas V. However, this solution may be not effective with large tanks, since required thickness of their walls quickly grows with size (so that the tank could withstand inner pressure). Tanks with frame structure are more simple to develop and their walls may be thin enough, since the loads are carried by stringers and the inner frame. Aluminum allows also may be used instead of steel for large balloon tanks. Titanium alloys are used as well. The Shuttle external tank has such structure, as well as most other large tanks. Small tanks often do not need framework. Intermediate solution, semi-monocoque, is

3 also applied, where both the frame and the skin carry loads
also applied, where both the frame and the skin carry loads. On the first rockets (like the V-2) the skin carried only hydrostatic loads of the liquid, while dynamic loads were carried by the outer frame structure. The same design was applied on the N1 Soviet lunar rocket. However, now this design is usually implemented only on small upper stages and on spacecraft. In most of launch vehicles, tanks carry all loads, including the weight of the upper stages. Tank assembly of a stage of a launch vehicle usually contains two separate tanks, for the Fuel and for the oxidizer. Often these tanks have own domes each one and are connected with an intertank structure that transfers loads between the tanks. However, sometimes there is a common bulkhead between the tanks, so the intertank is eliminated to save weight. Such design had the second stage S-II of the Saturn V. If at least one of the propellant components is cryogenic, specially if LH2 is used, extensive thermal insulation is required. The common bulkhead should also be insulated to avoid freezing of one of the components. In the inner volume of tanks slosh & vortex baffles are often present which dampen liquid sloshing and prevent vortex formation. Propellant sensors are placed to control the loaded and consumed amount of the propellant. Suction lines from the upper tank go to the engines in the bottom of the stage; they may pass through the lower tank (as on the first stage S-IC of the Saturn V) or along the external surface of the stage (as on the S-II or on the Shuttle external tank). The tanks of the upper stages and of spacecraft are constructed considering the requirement to reduce the surface (and so the weight) of their walls. This is the reason why they often are spherical. It is also important to keep the stage (and the spacecraft) compact at launch to decrease the size and the weight of interstages and of the nose fairing. Thus toroidal tanks are built, which surround the spherical tank of the second component, the engine or other assemblies. The example is the Blok D/DM upper stage of the Proton/Zenit launch vehicle. Of course, similar design is impossible on the lower stages to avoid very large

4 diameters (which would lead to high drag losses)
diameters (which would lead to high drag losses). However, if the volume of one component is significantly smaller than that of another one, its tank may have flattened shape, as in the case of the LOX tank on the S-II. If the engine of a rocket stage or a spacecraft should be fired in space, a problem of engine start in weightlessness appears. Liquid flies freely inside the tanks and does not cover the inlets of the suction lines. It also may be mixed with gas, spume may be present. So if the engine is fired, gas may get into the inlets and the engine may stall. To separate the liquid from gas and to direct it towards the inlets various methods are applied. One of the most frequent methods is to create non-zero weight for a short period of time. The rocket should be accelerated for that. On stages of launch vehicles it is done with special ullage motors (see further), on spacecraft where engines should be restarted many times this task is often performed by vernier motors or attitude control motors. However, the same problem is present in the smaller tanks of verniers. In small tanks special membranes are used to separate the gas from the liquid, and pressurizing the gas section above the membrane forces the propellant to the inlets. The pressurizing gas is held in special vessels or is produced by a gas generator. Fuelling The amount of propellants fuelled into the tanks is never equal to the value that follows from theoretical calculations. One of the reasons is technical tolerances. The exact parameters of the engines are never known, so in flight propellant consumption rate may be higher than demonstrated at static firings. In addition, some problems may arise in flight, for instance, earlier cut-off of one of the engines or other failures. The precise path of the rocket is also unknown at the time of the lift-off, since several deviations from the nominal trajectory may

5 occur; and each trajectory needs its own characteristic velocity
occur; and each trajectory needs its own characteristic velocity. Thus, in the real flight an uncertainty exists in the amount of propellant which should be consumed to deliver the payload to the target. By that reasons, some additional amount of propellant should be fuelled to compensate these random disturbances. This is the propellant reserve. Of course, it is impossible to fuel much more propellant than needed for a normal flight, since the propellant reserve is not normally used; actually it is a part of the structure mass and affect the mass ratio. For an upper stage, every extra gram of propellant diminishes the mass of the payload by one gram. For lower stages the impact is smaller, but it also plays role. Thus, the amount of extra propellant loaded is determined by soluving an optimization problem: the reserve should be sufficient to guarantee accomplishment of the mission task when the parameters are inside the tolerances and also in the case of several technical problems resulting in excessive propellant consumption, but at the same time not to significantly impact the payload. For the Saturn V, the criterion was % assurance that the launch vehicle will completes its primary mission. The amount of extra propellant is ordinarily about 1% – 2%. To diminish the mass of the propellant reserve, several methods may be used. First consider that the mass of the loaded oxidizer is usually much larger than that of the fuel. That means that a fuel excess, having moderate mass, enables to consume much more oxidizer. This is why a little bit more fuel is loaded than the nominal mixture ratio requires. A more radical solution is the propellant consumption control system. This system tracks in a real time the amounts of both components and slightly shifts their expenditure ratio so that by the end of the burn there were no excess of one component when another is

6 Stages and their separation
depleted. This system enables to consume all the propellant. This solution is particularly effective if close-loop navigation is used. In this case there is no need to cut off the engines of the stage when the predefined velocity is achieved. The engines may work until total depletion of the propellant whatever amount is loaded, thus the reserve will also be consumed. This means that this consumed reserve is the reserve also for the upper stages. So there is no need to have dedicated reserve on each stage, the reserves of all stages form the common reserve of the rocket, and the efficiency increases. Besides the reserve, some amount of the propellant remains unburnt being trapped in the suction lines, engines, on the bottom of the tanks and due to tolerances of the propellant sensors. Stages and their separation As we have already seen, there are two principle designs of a multi-stage rocket: serial (or tandem) staging and parallel staging. Most of launch vehicles include serial stages, but parallel stages is also frequent. Serial and parallel staging Parallel stages usually represent strap-on boosters (with liquid-propellant or, more often, solid propellant engines) which are ignited on the ground together with the main (second) stage which continues working after separation of the strap-on boosters (however, some of the boosters may be ignited in flight, like 3 of 9 boosters of the Delta II rocket in the 9-booster configuration). The principle task of the boosters is to provide additional thrust at the lift-off.

7 The first launch vehicle to implement such design was the R-7, which after further upgrades became the Soyuz and the Molniya launch vehicles. The main reason for such construction was possibility to ignite all engines on the ground and thus to avoid engine ignition in flight: this operation was considered problematic. Simultaneous ignition also allowed to check if all engines work normally, and if a problem occurred, it was possible to cut off the engines and to cancel the launch. Today if parallel staging is used, engines of all parallel stages are generally ignited at the same time by similar reason. It is specially important for the Space Shuttle: first the main engines are ignited and work for some seconds so that their safe functioning is made sure; later the boosters are ignited. However, a stage working at the sea level and in denser atmosphere looses the specific impulse and the thrust, so the parallel staging is in the general case less effective than the serial staging, where upper stages work in rarefied atmosphere. This is the reason why on the Titan IV the liquid core stage was ignited in flight, shortly before the SRB separation (contrary to the previous Titans, where the core stage was ignited on the ground). Serial staging is usual on the second and upper stages. Many rockets have only serial staging, without parallel boosters. A lot of projects of propellant transfusion from parallel stages (boosters) to the core stage have been developed. Such design may improve efficiency, since parallel tanks are emptied and may be jettisoned earlier. So they may be accelerated to a lower speed, thus more energy of the propellant is spent on acceleration of the upper stages. However, this design is more complicated and less reliable, additionally it requires identical propellants on the boosters and the core stages, so it have not still been implemented.

8 Separation of stages When a stage have provided its payload with the required velocity increment or have worked until the depletion of the propellant, it should be separated. Stages are connected so that their connections may be severed by pyrotechnic devices. In the case of strap-on boosters these are explosive bolts and/or frangible nuts. They hold the stages together until the electric signal for separation. By this signal detonation occurs, the bolts/nuts break apart and the structural connection between stages vanishes. The structural tie between serial stages is broken by an explosive device around the circumference of the stages which severs a tension strap. When the structural tie between the stages is removed, it is important to quickly withdraw the spent stage from the vehicle. The reason is that due to the impulse of aftereffect the spent stage may continue moving ahead and slam into the payload damaging it. Several methods are applied for quick withdrawal of the spent stage. In the case of a strap-on boosters separation, the engine of the core stage works in the moment of separation, so the vehicle moves with acceleration relative to the separated boosters. The problem of igniting the upper stage is thus non-existent. So the problem of separation is reduced to carrying the boosters apart from the vehicle. This is performed with the aid of special separation sequences which eliminate the danger of collision between the stages during the separation process. Often special small engines are provided that carry the separated stage away from the vehicle or, alternatively, their role may be played by damp of gases from the separated stage. On the Soyuz, pressurization gases are used, on the Shuttle

9 SRB, assemblies of four solid rocket motors are present on each end of the SRB.
On serial stages separation is accomplished by the fact that the engine of the upper stage does not work and the upper stage has no initial acceleration relative to the stage to be separated. Two principle separation patterns may be pointed out, these are fire in the hole (or hot separation) and cold separation. Fire in the hole is the separation method when the engine of the upper stage is ignited while the stage is still attached to the lower stage to be separated. After the ignition, the connections between the stages are severed and the lower stage is repelled by the exhaust gases. This method enables to solve a number of problems. First, the problem of ignition of the upper stage is generally solved: the ignition is performed when the rocket moves with acceleration due to the impulse of aftereffect of the lower stage, and the propellant is settled on the bottom of the tanks. Second, the spent stage is effectively pushed away by the exhaust. However, special measures should be taken that the exhaust do not damage the upper stage. An escape holes should be provided to the exhaust gases between the stage. To let the gases escape, the interstage has a lattice structure with large holes, as it is done on the Proton (1st stage/2nd stage interstage) and the Soyuz (2nd stage/3rd stage interstage). In addition, the upper side of the lower stage should be protected by ablation from excessive thermal loads, so that the hot gases do not reflect back and not damage the upper stage and that the tanks of the lower stage do not burn through and explode at separation. If cold separation is performed, the engines of the upper stage are ignited only after full separation of the lower stage. This method enables to decrease loads during separation which

10 occurs in more gentle conditions
occurs in more gentle conditions. However, the problem of the spent stage withdrawal arises. The spent stage may be repelled by thrust of pressurization gases escaping from valves which are opened on the separated stage. It the case of a solid stage, additional nozzles may be opened by explosive devices in the body of the stage which provide thrust in the opposite direction. Retrorockets are another way to quickly decelerate the spent stage. These are solid rocket motors installed on the lower stage which are ignited at the separation and provide brief but high retrograde thrust. Retrorockets are used to withdraw the second and the third stages of the Proton, they were also used on the first stage of the Saturn 1/1B/V. Another problem to be solved is engine ignition. After separation the only noticeable force influencing the vehicle is gravity; this means that the vehicle is in weightlessness. In these conditions nothing forces the propellant to the inlets of the suction lines, thus suction may be ineffective. In smaller tanks the problem is greater, specially if the engine should be ignited after a lengthy period of coast: in this case the propellant may even mix with the pressurizing gas and the inlets will be exposed to the gas. This issue may be solved in different ways. In small tanks membranes separating gas from liquid are frequently installed, and the propellant is forced to the inlets by pressure. Other method is to initially accelerate the rocket. This may be done by special ullage motors: these are small motors, often solid, that are fired for several seconds to settle down the propellant. Such motors were used, for example, on the 2nd stage of the Saturn V (on the last Saturns ullage motors were deleted). The required acceleration may be provided by other motors as well. On the 3rd stage of the Proton the vernier engines are ignited before separation and serve as ullage motors. On the Apollo Lunar Module ullage was provided in space by its verniers as well. On the 3rd stage of the Saturn V the Auxiliary Propulsion System

11 was installed to control its attitude at docking, this system was also used for settling the propellant. Interstages have ordinarily cylindric shape and are separated together with the stages. Constructors usually try to decrease their mass since it directly impacts the mass ratio of a stage. A less massive interstage should be shorter (shorter cylinder could also be thinner and lighter to withstand the loads; long cylinders loose stability undr smaller loads and thus should have thicker walls). However, the interstage should be long enough to accommodate the nozzle of the upper stage. The nozzle may be quite long on stages working in vacuum to allow higher expansion ratio. So, the requirement of high expansion ratio contradicts the requirement of short interstage, and sometimes this contradiction is resolved by making an extendable nozzle. At the lift-off the nozzle extension is in the stored position and after the separation it is extended. Stages of launch vehicles often include a Range Safety Sysem – that is the system that ensures destruction of the rocket or its part in the case of a launch failure if the falling debris jeopardize populated areas. The system is based on explosive devices that, if detonated, destruct the rocket body into small parts and disperse the propellant to avoid massive fires on the place of fall. The system is activated by a command from the ground if the rocket wanders off its trajectory beyond the allowed limits. The Shuttle and other US vehicles are also provided with such system since launches occur not far from densely populated areas.

12 Nose fairings Most of payloads of launch vehicles should be protected at launch from aerodynamic, thermal and acoustic (sometimes also g-force) loads, since during accent the nose part of the rocket is subject to strong aerodynamic forces and high temperatures. To protect the payload, a nose fairing (also called nose cone, or nose shroud) is installed. At lift-off, the payload is covered (fully or partially) by the fairing. The requirements that the fairing should meet are as follows. It should be sufficiently large so that the payload can be accommodated inside it. At the same time its size (mainly width) should not be excessive in order not to increase drag losses. Since these two requirements are conflicting, different sizes of fairings are often offered on the same launch vehicle for different payloads. For example, Delta IV Medium goes with two fairings, one with the diameter of 4 m and another with the diameter of 5 m. The latter variant delivers larger payloads but has smaller mass capacity for GTO missions in spite of a larger third stage; the reason is increased drag. (However, boosters may be added to increase the mass capacity). Titan fairings may be used as well. To decrease mass of the fairing, it is sometimes made from composite fiber-reinforced plastic. Being more expensive, such solution increases the payload. Nose fairings are useful mostly only in the dense atmospheric layers. That the rocket ascends higher, aerodynamic loads drop off quickly, and the fairing becomes useless and only worsens the mass ratio. So it is usually jettisoned at a proper altitude (often soon after the first stage separation). The jettison should be done accurately, so that the fairing or its parts do not hit the vehicle. So fairings usually consist of two or more petals tired together at launch. At the jettison, the tires are cut by pyrotechnic devices and separated from the vehicle flying

13 apart. However, in rare cases the fairing is not separated intentionally during the ascent and is jettisoned only on orbit. This was the case of the very first satellite PS-1: the fairing was jettisoned in weightlessness for the reasons of simplicity. During the launch of the US space station Skylab the fairing was not separated for another reason. At launch it played a role of the bearing structure for the astronomical unit which was connected with the body of the station with a foldable light truss that could not have beared the g-loads of the launch. In some cases a compact upper stage of a launch vehicle is also covered by the fairing. This enables to diminish the mass of the upper stage, for example, by deleting the outer protective shell. The overall payload of the rocket increases, since the fairing, although being larger in this case, is jettisoned early, but a lighter structure of the upper stage should be delivered to the target orbit. In some cases the payload is covered with the fairing but partially. The Apollo spacecraft had only minor protective cover above the front side of the Command Module, this cover went as a single unit with the Launch Escape System and was jettisoned together with it. Nose fairings may sometimes have holes in their body for servicing the payload before the launch: fuelling, checks etc. These holes are closed by hatches before the launch. In the case of crewed spacecraft, the crew enters through a special hatch (on the Vostok and the Gemini, where the crew was provided with ejection seats, astronauts should be ejected through this hatch if a launch failure occurred). In the case of a failure on the Soyuz manned spacecraft, two modules of the spacecraft are carried away from the vehicle together with the upper section of the fairing which holds them at the emergency escape until the spacecraft is at a safe distance from the vehicle; when the fairing is jettisoned.

14 Launch escape system Most of expendable launch vehicles designated to deliver crew into orbit have been provided with a launch escape system (LES). This is a special rocket device attached to the manned spacecraft. The task of this device is to quickly separate the spacecraft from the launch vehicle and to carry it away to a safe distance in case of emergency. In all designs that have been applied so far LES have had similar constructions. They consist of a spire-shape solid rocket motor placed atop the vehicle and attached to the spacecraft directly or through the nose fairing. The motor has a system of nozzles directed back by some angle to the axis of the rocket (so that the spacecraft were not damaged by the exhaust gases). In case of emergency the entire spacecraft or its crew module is separated from the rest of the vehicle and the solid rocket motor is fired for a short time. To quickly withdraw the module from the vehicle, high accelerations should be provided, so the crewexperiences high g-loads (up to 15 g and higher), but only for some seconds, so no serious injuries should occur. The period of the firing is about 10 sec, it is defined by the need to raise the crew module to a sufficient height if the failure occurs on the launch pad (so that the parachutes may open safely to soft-land the module). Besides the solid rocket motor, the LES includes some other systems, among which are aerodynamic control surfaces that stabilize and/or orient the module after the launch escape sequence is initiated, a motor that sends the module apart of the local vertical line (pitch control motor) if the failure occurs on the launch part, a LES jettison motor, electronic equipment etc. The command to start the launch escape sequence is sent either by the ground control (only in case of a launch pad emergency) or by electronics of the launch vehicle when it

15 detects anomalies. The LES may be activated only until the vehicle reaches a certain altitude, where the LES is jettisoned together with the nose fairing (on the Saturn V it was jettisoned soon after ignition of the 2nd stage). If dangerous situation develops later in flight, the spacecraft may be withdrawn by its own engine, as on the Apollo. There are alternative developments that have not been applied in practice so far; the LES may be placed below the spacecraft and push it from tail, not to pull it by nose. In a normal launch this system may be even used for a final orbital insertion maneuver. Not all manned spacecraft have been provided with a LES. First it appeared on the Mercury spacecraft. The Vostok & the Voskhod did not have any LES, but the Vostok was provided with an ejection seat that should have ejected the cosmonaut in case of emergency at low altitudes (at altitudes higher than 4 km the engines of the launch vehicle should have been cut off and the spacecraft should have been separated by a nominal sequence). The ejection seat of the Vostok were used also at landing, it ejected the cosmonaut from the capsule, so he could land using his own parachute (this was done since the spacecraft has too high speed at parachute-landing). Analogically, there were no LES on the Gemini spacecraft, the astronauts were provided with ejections seats like on the Vostok (but did not use them at a nominal landing). On the Voskhod the ejections seats were deleted. On the Space Shuttle the ejection seats were installed on the first four test flights, when the crew consisted of two men. Later the ejection seats were deleted since it is impossible to proved the whole crew of 7-8 astronauts with such seats. After the Challenger disaster, a special equipment was installed that lets the crew to leave the vehicle one by one through the main hatch and to parachute-land, but such evacuation is possible only in a steady glide. The Shuttle also has a number of abort modes in case of several launch malfunctions, which include an emergency separation of the orbiter and landing on the Canaveral or on one of the emergency bases on the Eastern

16 coast of the Atlantics, or insertion into a lower orbit
coast of the Atlantics, or insertion into a lower orbit. But in general the crew of the Shuttle is not protected from such catastrophic events as the Challenger disaster. Launch emergency escape have saved two crews. The first accident occurred on Apr 1975, when the Soyuz launch vehicle with the Soyuz 18-1 (unofficial designation) spacecraft failed due to the 2nd and the 3rd stages separation malfunction. The LES had been already jettisoned by that moment, and the spacecraft separated from the launch vehicle following the launch escape sequence. Due to unpredicted behavior of the descent maneuvering system the crew experienced very high g-loads (briefly up to ~20 g), but landed with no serious injuries on the USSR territory and was rescued. The second accident occurred in Sep 1983 with the crew of the Soyuz T-10-1 (unofficial designation). The launch vehicle caught fire on the launch pad due to a malfunction of one of the engines. The launch escape sequence was initiated by ground officers several seconds before the booster exploded, and the LES carried away the spacecraft from the launch pad in fire. The spacecraft safe-landed several kilometers from the launch pad.

17 End of the Lecture 13

18 Saturn V first and second stages
S-IC first stage. Intertank structure is seen, suction lines inside the kerosene tank, slosh baffles (By source) S-II second stage. Common bulkhead is seen, suction lines outside the LOX tank, ullage rockets (By source)

19 Blok DM-SL Upper stage for Zenit-3SLB. The spherical LOX tank and the toroidal kerosene tank are clearly visible (By source)

20 Rocket staging (By source)

21 Separation of strap-on boosters
Upper link Separation motors Bottom link Engines of the stages (By source)

22 1st stage with the interstage of the Proton (By source)
Lattice interstages Soyuz assembly, 1st stage/2nd stage lattice interstage is seen (By source) 1st stage with the interstage of the Proton (By source)

23 Extendable nozzle of RL-10B-2 engine
Extendable nozzle of RL-10B-2 (Delta III/IV) in the stored position (By source) Nozzle of RL-10B-2 in the extended position (By source)

24 Akari IR observatory inside its nose fairing atop M-V (By source)
Nose fairings 5 THEMIS satellites for aurora studies together with a PAM inside the nose fairing of their Delta II (By source) Akari IR observatory inside its nose fairing atop M-V (By source)

25 LES of the Apollo spacecraft (By source)
Launch Escape System LES of the Apollo spacecraft (By source) LES of the Soyuz spacecraft (By source)


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