introduction to astronautics sissejuhatus kosmonautikasse n.
Skip this Video
Loading SlideShow in 5 Seconds..
Introduction to Astronautics Sissejuhatus kosmonautikasse PowerPoint Presentation
Download Presentation
Introduction to Astronautics Sissejuhatus kosmonautikasse

Introduction to Astronautics Sissejuhatus kosmonautikasse

283 Vues Download Presentation
Télécharger la présentation

Introduction to Astronautics Sissejuhatus kosmonautikasse

- - - - - - - - - - - - - - - - - - - - - - - - - - - E N D - - - - - - - - - - - - - - - - - - - - - - - - - - -
Presentation Transcript

  1. Tallinn University of Technology Introduction to AstronauticsSissejuhatus kosmonautikasse Vladislav Pustõnski 2009

  2. In astronautics the task to return a spacecraft or its part from space or to land a space probe on a planet is common. This task arises when a manned spaceship returns to the Earth with its crew, when a space probe parachutes in the atmosphere of a planet transmitting data or a probe soft-lands on a planet delivering scientific payload to its surface. Sample return missions should soft-land on the target body and than to perform landing on the Earth. In this lecture we will concentrate on methods of descent and landing on a celestial body. Descent and landing Landing from orbit and direct landing Landing from orbit takes place when a spacecraft before landing is placed on a (low) orbit around the celestial body. This is mostly the case of manned spaceships (except for the lunar Apollo missions) and many planetary landers. Direct landing is the case of a spacecraft approaching a celestial body from deep space (or a very high orbit) and performing landing directly from the trajectory of approach. If the spacecraft is initially situated on a low Earth orbit, it descends from this orbit and lands. But for a deep space probe a proper landing method should be chosen: either to land directly or first to go into a low orbit and to land from this intermediate orbit. Both methods have advantages and drawbacks. Direct landing is more simple, since many operations related to orbital insertion are skipped. For orbital insertion, precise navigation is needed (this maneuver is usually performed automatically without ground control intervention because of a large distance), the parameters of the orbit should be ensured with high precision. In addition, the engine of the

  3. space probe should restartable, and a longer in-orbit lifespan should be provided. Very important is also the fact that orbital insertion requires expenditure of propellant. This expenditure is unavoidable in the case of a celestial body without atmosphere, but if the planet has atmosphere, no propellant is required in the case for a direct landing: entry to the atmosphere would cancel the initial hyperbolic velocity. Propellant economy is somehow balanced by the need for a heavy heat shield and a more robust lander design capable to withstand atmospheric entry at hypersonic higher velocity. Drawbacks of direct landing include impossibility to select a better landing site on poorly studied planet. The landing area is defined by the trajectory of approach. It may be shifted during the flight, but observations from distance are not effective to choose a suitable landing area. Detailed images of the landing area may be obtained only on the final phases of approach, when the landing point cannot be effectively shifted. Landing from orbit allows preliminary close-range observations of the surface and selection of a better (safer or more interesting) landing site. This is the reason why the Viking Martian probes first went into orbits and only after having performed observations of the surface suitable landing areas were selected and the landers were released to soft-land in these areas. Contrary to the Vikings, the first lunar landers (the Luna-9/13, the Surveyors) landed directly from the trajectory of approach: this simplified the guidance (especially in the case of the Lunas), and the probes also did not carry orbital modules that could have photographed the surface with anticipation. The Martian landers (the Mars-3/4/6, only the first is a partial success) also performed direct landings due to guidance and weight economy reasons. The Jupiter atmospheric probe carried by the Galileo Jupiter orbiter was separated before the orbital insertion of the probe, it

  4. allowed not to brake down the probe and to safe weight. If the probe did not separate, the orbiter would have failed to go into orbit. Vice versa, the Titan atmospheric and soft-landing probe Huygens was inserted into the Saturn orbit jointly with its carrier probe Cassini. This allowed to send the probe to Titan with higher precision and to provide the right entry angle (the entry angle was of higher importance for the Huygens since it should have followed precisely its predefined path in the atmosphere, contrarily to the Jovian probe of the Galileo that could withstand extreme g-loads). Nowadays Martian landers usually perform a direct landing since guidance have improved dramatically since the early era of Mars research with space probes, and better maps of the surface are available now as well. So today it is possible to select the landing site before the launch of the probe and to aim it to the desired area. As for the landing precision, several factors play role. In the case of a direct landing, the descent trajectory is shorter and the descent time is shorter as well. So initial deviations from the nominal path cause smaller shifts from the nominal landing point than it takes place at landing from orbit, when initial deviations may cause larger landing point dispersion. On the other hand, orbital insertion allows to adjust the trajectory and the time of starting the landing sequence, so the descent trajectory can be set with higher precision. In addition, due to longer descent path and time there are more opportunities to make trajectory adjustments. So, the landing precision in this or that case depends on ability to set precisely the approach trajectory and on descent guidance precision. For the lunar sample return missions Luna-16/20/24 the correct landing place was critical (in order to guarantee sample return), so they first went into lunar orbit and adjusted their trajectories for several days before they were ready for a final descent. Gravity field of the Moon was studied and mascons were taken into account. On the

  5. contrary, Apollo CMs which returned from the Moon landed with higher precision when Apollo CMS which landed from LEO. Generally, the exact landing point is unknown, and the predictions are represented in the form of landing ellips. The center of the ellipse corresponds to the point of highest landing probability, the probability decreases towards the edges of the ellipse. The edge usually corresponds to the probability estimation of 3 (99.7%). Further we will discuss soft-landing methods, which imply landing (or descent) with moderate final velocity. So hard-landing (landing with orbital or nearly orbital velocity, which mostly mean crash-landing) will not be considered. Some earlier lunar probes performed intentional crash-landings, like the Luna 2 (the principle task of which was to achieve the surface) and the Rangers (they radioed close-range photos of the lunar surface but were not intended for soft-lading). However, hard-landing does not imply crash-landing; for instance, penetrators may hit the surface at a high speed but be designed to withstand the impact and to penetrate the surface in order to send scientific data from underground surface layers. Penetrators were aboard the Mars-96 and Mars Polar Lander Martian probes, but both probes failed to complete their mission. Soft-landing implies canceling the initial speed of the probe, either orbital or hyperbolic. Canceling methods depend on the celestial body where the spacecraft lands. There are two different types of landing: landing to a body with atmosphere and without atmosphere. Further we will discuss both.

  6. Landing on a celestial body with atmosphere If there is atmosphere around the celestial body (we will call it planet, although it may be Titan – the only satellite with dense atmosphere), it may by used for canceling the velocity. In most cases there is no sense to miss this opportunity, since it allows to significantly reduce the amount of the propellant onboard and weight of the engines thus gaining the mass of payload. However, engine cannot be fully omitted. It is needed for corrections of the approaching trajectory and/or for adjusting maneuvers on the intermediate orbit. Once the spacecraft is on the orbit, the (re-)entry maneuver is needed: the orbital velocity should be reduced to place the periapsis inside the dense atmospheric layers (or even below the surface). It is a so-called deorbit burn. The first manned spaceships had special retrorockets which provided the reentry impulse. These retrorockets had to be very reliable since safe return of the astronaut depended on their work. The Vostok spacecraft had no backup engine and thus they were put into orbits with a very short decay time. If the retrorocket engine failed, the spacecraft could reenter before the resources of the life support system would be depleted (the Gagarin’s spaceship was put on a higher orbit because of a minor failure of the launch vehicle, but the retrorocket engine provided safe return). On contemporary spaceships main engines perform the reentry impulse and they are always backuped. For reentry, a spacecraft turns its tail forward and fires the engine. Actually, the reentry impulse is directed a little bit down, under some angle with the orbital velocity. This is done to increase the reentry angle. Typical characteristic velocities of reentry impulse are about 150-200 m/s for reentry from LEO.

  7. Entry angle, g-loads and energy dissipation Spacecraft which enters the atmosphere of a planet possesev very high kinetic energy. Canceling orbital velocity actually implies canceling the kinetic energy of the spacecraft by converting it into heat. Specific kinetic energy (per unit mass) may be estimated as V2/2 (V being the velocity at the entry). For reentry from LEO it is ~30 MJ/kg, for return from the Moon or deep space it is ~65 MJ/kg. For a direct entry into Venus atmosphere it is even larger, for a direct entry into Martian atmosphere it is smaller. However, heat of fusion of construction materials is much smaller: for example, for aluminum it is ~0.4 MJ/kg. This means that it is impossible to convert the kinetic energy of the spacecraft to heat of its parts apart of having a very heavy heat shield. So, the energy should be dissipated by heating the air: this is the basic principle of air braking. Two principle requirements should be met during an atmospheric descent. First, the amount of heat absorbed by the spacecraft should not damage it. Second, g-loads during deceleration should not exceed safe values. g-loads are defined by deceleration time: if the entry speed is fixed, g-loads are smaller for longer time of braking. So, entry trajectory should be longer to reduce g-loads, thus entry angle should be smaller. Slower deceleration rates also reduce intensity of energy dissipation, and instant heat loads are smaller. On the other hand, total amount of absorbed heat increases with the length of the trajectory since the exposition time to the heat flux rises. Entry angles cannot be very small for a direct landing: too small entry angle may cause insufficient braking and the capsule may not be captured by the atmosphere and will “jump out” back to space. It may than remain on an elliptic orbit and reenter again on the next orbit or, if deceleration was very moderate, it will still have a

  8. hyperbolic excess and will leave the planet. The first Venus landers entered under an angle of 620 – 650, so they were subject to extreme g-forces which might achieve 450 g. Manned spaceship are limited by g-loads of ~12 g (the case of the first Vostok/Voskhod and Mercury capsules; g-loads for the Shuttles do not exceed ~2 g). The double entry method (so-called skip reentry) may be used intentionally for shifting the landing point and for reducing the heat loads. During the first entry the speed is partially canceled and the capsule returns into space for cooling down. On the next orbit it reenters and performs the final landing. This method was used by Soviet lunar fly-by probes Zond-6/7: double entry allowed them to land on the territory of the USSR. When a capsule enters atmosphere with a hypersonic speed, a shockwave is formed in its front. It moves together with the capsule, temperature of air compressed in the shockwave may rise to thousands of Kelvin. Energy dissipates mostly by radiation of the heated air (at speeds ~6 km/s and higher, at speeds higher than 8 km/s radiation dominates) and by convection in the boundary layer. Heat flux spreads in all directions, as well as towards the capsule. Large entry angles mean quick growth of the flux and its quick decline, smaller angles cause gradual growth and decline of the flux (and g-loads). However, since air density drops by a nonlinear (actually, close to exponential) law, the dependence on the entry angle is ambiguous. In some cases a small entry angle may cause slow initial deceleration in the upper layers with subsequent dive into dense layers and shock braking. During the phase of rapid deceleration the spacecraft is surrounded by a cloud of plasma which makes impossible communications with the capsule: this is communications blackout. The communications restore after the rapid braking accompanied with air ionization is over.

  9. To protect a capsule from extreme heat fluxes and erosion, expendable thermal protections are used, among which ablation and heat sink structures. For thermal protection, the capsule is provided with a heat shield which is more thick in the front side and is thinner at the back side where heat fluxes are smaller. The working principle of ablative protection is that the exposed surface is covered with special material that sublimes; phase transition is capable to absorb a lot of heat. The heat is carried away by the gases. These gases also form protective layer between the hot shockwave and the heat shield. They are opaque and thus block radiative heat transfer to the capsule. Opacity is provided by carbon particles which are formed in the ablative layer by coking. Heat sink is provided by a material with high specific heat, for instance by metals. Heat flux that passed through the ablation is absorbed by this material and do not pass further to the structure of the capsule, later it is irradiated back into the atmosphere. This implies heavier heat shield designs. The heat shield is separated from the main structure of the capsule by material with low heat transfer coefficient. When the capsule has lost its speed and heat loads have dropped off, the hot heat shield is often jettisoned: this lessen the weight of the capsule, so smaller chutes are required, and also this prevents further passage of heat from the shield to the structure. The mass of the heat shield achieves ~1/3 of the total mass of the capsule (for Vostok it was ~800 kg, the total mass of the capsule being ~2500 kg). Space Shuttles, as well as their Soviet analog Buran, have reusable thermal protection. It consists of insulation tiles of refractory materials (ceramics, fibrous composites etc.) which possesses very low heat conductivity and specific heat. In different areas different tiles are used, depending on temperatures and heat fluxes. Reinforced carbon-carbon tiles can

  10. Shapes of entry capsules withstand temperatures higher than 1300 0C. Density of these tiles is very low, so the mass of the heat shield is moderate. Its working principle is insulation, not ablation nor heat sink. In each flight some tiles are lost, but this damage is not critical for the vehicle. However, the disaster of Columbia in 2003 occurred due to a fatal destruction of tiles in a leading edge of the left wing. The tiles were damaged by a piece of insulation foam which fell from the external tank: tiles are very fragile. The simplest shape of entry capsule is spherical; such were the capsules of Vostok/Voskhod and of Venera Venus landers. A spherical capsule do not provide any lift force (so its lift-to-drag ratio, or L/D ratio, is zero) and is capable only for a ballistic descent. To avoid chaotic spinning of the descending capsule, its center of mass is placed in front of the center of pressure. Thus g-loads are directed along one axis; this is particularly important for a manned spaceship, since human body withstands much better g-loads in front/back direction. More often entry capsules have a shape of blunt body. A blunt body may possess a lift force which depends on its angle of attack in the flow. The lift force may be used to increase the path length and so to decrease peak accelerations and g-loads. It is also possible to actively maneuver in the atmosphere by controlling the lift force and thus to increase the precision of landing. Different shapes of blunt body capsules are applied. On manned spaceship the forebody is usually spherical, and the afterbody is conical (Apollo, Gemini, Mercury) or bi-conical (Soyuz,Shenzhou). Thermal protection in the forebody is more extensive since it is subject to the largest thermal loads. Lateral and afterbody are shielded in smaller extent. The

  11. center of pressure is placed on the symmetry axis and the center of mass is in front of it, shifted slightly from the axis by corresponding positioning of heavy equipment. This ensures the position in the flow under a certain pitch angle which defines the L/D ratio (for the Apollo Command Module it was about 0.37). Lift control is usually performed by roll of the capsule with special engines (other engines adjust the pitch angle if needed). Thus the direction of the lift force changes from vertical to lateral, so the capsule may not only to change the vertical velocity but also to perform side maneuvers. Attitude control is performed by the onboard computer. The Apollo CM before the peak g-loads (of about 6 g) maneuvered to reduce the g-loads, than it adjusted the landing point so that landing precision of ~20 km was achieved. The flight path angle at the entry (about 70) was set by orbital corrections. If controlled descent is impossible due to some software or navigation error, the capsule performs ballistic descent. In this case it is spun by roll engines so that the lift force takes arbitrary values. This is done to prevent accidental negative lift force with extreme g-loads. Such situation happened during the emergency descent of the Soyuz-10-1, when the launch vehicle failed and the crew had to withstand peak g-loads of ~20 g. Dispersion of landing points during ballistic descent is larger, but ballistic descent allows a safe return. For the Mercury it was the only available descent method, as well as for Vostok/Voskhod. Some entry capsules have a sphere-conic or cone-conic shape, where the forebody has a shape of cone. A cone, specially with small half-angle and center of mass shifted forward, is more stable in the flow, so this design is mostly applied on capsules that are not actively steered during the entry. Aeroshells of planetary landers are often conic. Winged design allows full control over the landing process and allows accurate

  12. Final descent maneuvers by range as well as lateral maneuvers: for example, the Shuttle is able to perform landing at distances up to ~2000 km aside from is orbital path. Wings allow to land with precision of tens of meters, as it is possible for aircraft. The highest drawback of this design is that the weight of wings is a substantial fraction of the total weight of spacecraft, and their only function is to enable landing; wings are useless in space. Another disadvantage is that only controlled descent is possible for a winged spacecraft: being aerodynamically unstable, it cannot perform a ballistic descent and would be destroyed if the control fails. So this design requires very high reliability. During the phase of intensive aerodynamic braking the largest fraction of the kinetic energy of the capsule is canceled, and it falls under the influence of gravity field with a speed which depends on its size and geometry. Near the surface of the Earth this speed is in the range of ~50 – 150 m/s. For a safe lading, the touchdown speed should be markedly smaller. If there is a crew onboard, landing on water requires ~ 12 – 15 m/s, on ground ~ 6 – 9 m/s. To reduce the final speed to these values, parachute systems are usually applied on Earth since they are very effective in the Earth conditions. Ordinarily a chute system includes several chutes which serve to different tasks. The main chute is extracted by a pilot chute, sometimes a staged system of drogue chutes is applied (they kill the residual horizontal velocity), and stabilization chutes may be used as well to stabilize the capsule in the flow before introducing the main chute. Sometimes more than one main chute is introduced for the reasons of backup (this was the case of the Apollo CM; in the

  13. expedition of Apollo 14 one of the 3 main chutes failed, but the descent was safe). Otherwise, a reserve chute is present and should be introduced if the main chute fails (in the flight of the Soyuz-1 both chutes failed to deploy and the cosmonaut died). In flight, chutes are placed in a special compartment of the backside of the capsule. This compartment is opened usually by a signal of the pressure sensor. However, a chute system may be not the most effective method on other planets from the point of view of weight. On the Venus the density of the atmosphere near the ground is so high (~0.1 of the water density), that chutes are not needed for the final descent. The Venera soft-landers used chutes only slowly descending in the upper atmosphere in order to study these layers. At the height of some tens of km the chute was jettisoned and the final descent was performed with the aid of the braking & stabilizing ring at the top of the capsule (from the Venera-9). One of the small atmospheric probes of the Pioneer Venus, which descended with no chutes, even survived the impact and was able to continue transmission from the surface. Vice versa, the density of the Martian atmosphere is very small, so for effective braking a very large and thus heavy chutes would be required. Chutes of reasonable size can brake a capsule only up to speeds of ~200 – 300 m/s, and they should be introduced into the flow at hypersonic velocities. Cascading drogue chutes of increasing surfaces may be required to avoid steep rise of loads. Alternatively a reefed chute may be applied, it fully opens after initial braking (used on the Mars landers, but also in the atmosphere of the Earth: on Apollo capsules and others). By this reason sometimes only the main chute is aided by retrorockets. At the end of parachuting, when the residual velocity is several tens or even 1 – 2 hundred

  14. Landing on a celestial body without atmosphere m/s, retrorockets are fired and kill the residual velocity. Such system was applied on the Mars , the Pathfinder & the MER landers, retrorockets were secured to the lines of the chute. On the Pathfinder & the MERs airbags were applied as well to cancel the residual speed (see further). On the Viking & Phoenix landers the chute was jettisoned on a certain altitude and rocket-powered descent was applied on the final phase. On the Earth the density of atmosphere is also sometimes insufficient for effective braking without significant increase of weight of the parachute system, since the entry capsule may be quite heavy (>5 tons in the case of the Apollo CM). By this reason in the US landing on water was usual (Mercury, Gemini, Apollo). The first Soviet manned spaceships Vostok landed without the cosmonaut, who had to eject from the capsule and to parachute separately. The following spaceships (Voskhod, Soyuz) applied retrorockets (fired by command of a -altimeter) to cancel the residual speed before the touchdown. Sometimes parachuting capsules are captured in air, as it was done with the capsules of US reconnaissance satellites Corona. If the celestial body do not possess any substantial air envelope (as the Moon, the Mercury, most of satellites), only a rocket-powered descent is possible. Rocket-powered descent may be also applied on bodies with atmospheres if air density is insufficient for effective parachuting: for several types of missions this may be the case of the Mars. The principles of a rocket-powered descent do not differ very much from that of a general rocket flight, so they have been analyzed in the previous lectures. The problem of the soft-

  15. landing may be divided into two phases: 1) canceling the initial velocity (orbital or hyperbolic); 2) final slowing down and landing with a safe terminal velocity. To perform these tasks, the landing spacecraft should possess an engine powerful enough to quickly kill the initial velocity (in order to reduce losses, as we have discussed) and it may also have vernier engines to cancel precisely the residual speed before the touchdown. Alternatively, the residual speed may be canceled by passive methods, including airbags and/or dampers. The spacecraft should also possess a navigation and guidance system (which may be in some cases quite simplistic, as it was in the case of the Luna-9/13; this system have been already described in one of the previous lectures). A notable addition to the navigation system is that the spacecraft should be able to determine its position (and, in the general case, velocity) in respect to the target body. Usually radio altimeters are instruments that do this job. Their working principle is analyzing of the radio signal sent towards and reflected from the surface of the celestial body. Basing on the time of delay and on the Doppler shift it is possible to determine the distance and velocities. Optical imaging (automatic comparison of photographs made with a known delay) and laser altimeters may be applied as well. Rocket-powered descent may be used on celestial bodies with rarefied atmosphere (mostly on the Mars) for a final descent phase, thus replacing the chute system. The examples are the Vikings and the Phoenix. Being more complicated and less reliable (since sophisticated rocket motors and guidance hardware & software are required), this method enables to perform landing more precisely and the hardware may be lighter: no larges chutes needed nor heavy damper system for extinguishing residual excess velocity (airbags or something similar). The

  16. Touchdown and canceling of terminal velocity drawback is that exhaust gases from the nozzles mix with the soil and change its properties, so analyses of the soil are interfered by this contamination. The nozzles of the Vikings and the Phoenix were directed aside from the vertical in order to reduce contamination, and possible presence of exhaust chemicals in the soil was taken into account when analyzing data of soil studies. The Vikings performed a free fall from the height of 3 m, so that their engines did not contaminate the surface from a close distance. Before the final touchdown the engine is usually cut off at some small height above the surface (by a signal of a -sensor or something like this) so that the escape gases reflected from the surface could not damage the spacecraft. In some cases the capsule immediately before the touchdown possesses a residual velocity which is still not safe. This velocity should be canceled. In the case of a rocket-powered descent the residual velocity is usually small, about 1-3 m/s (except for simplified landing schemes), so damping mechanisms in the landing gear are sufficient to kill it. So the Surveyors, the Apollos, the Luna-E8 probes had only dampers in their landing gear. Parachute landing of a heavy capsule on the Earth or on the Mars are usually related with higher residual velocities, ~10 m/s or more. If the capsule lands on the ground, retrorockets are frequently used immediately prior to the touchdown, their task is to zero the terminal velocity. They are applied on Soyuz reentry module and were applied on the Mars, the

  17. Pathfinder & the MER capsules. All US expendable manned spaceships landed on water, since such landing allows residual velocities about 10 m/s and even more. At the splashdown the capsule enters the water under a certain angle to reduce g-loads. Sometimes retrorockets are not sufficient, so additional damping may be used, for example, airbags that cover the capsule before the touchdown. Airbags are inflated by gas produced by a special gas generator and damp the loads of the impact. The capsule jumps along the surface for a while until stops in a still position. When the airbags are deflated and retracted (MERs) or jettisoned (Luna-E6). Luna 9 was the first capsule provided with airbags. These were two inflated hemispheres which had been inflated before the landing sequence was started and surrounded the oval capsule forming a kind of a ball. They dampened the impact at the speed of >10 m/s and when were jettisoned. The capsule was verticalized by four petals which deployed and opened the capsule. A similar system was used on the Pathfinder and the MERs, where there were a number of spherical airbags which were inflated by a gas generator and protected the capsule which jumped over the surface after the chute was separated at the height of several meters. When the capsule stopped, the airbags were deflated and retracted, the capsule was verticalized by opening petals.

  18. End of the Lecture 16

  19. Skip reentry Skip reentry scheme (By source) Zond lunar fly-by probe (By source)

  20. Space Shuttle thermal protection Silicia tile from Atlantis (By source) Tiles on the wing of Shuttle. Black ones are newly replaced (By source)

  21. Shapes of entry capsules Landing capsules of the Voskhods (By source) Jupiter atmospheric probe of the Galileo (By source) Landing capsule of Soyuz (By source) Shuttle orbiter (By source)

  22. Manned entry capsules 1 – center of pressure; 2 – center of mass; 3 – heavy equipment (By source) Apollo Command Module (By source)

  23. CEV Orion parachute assembly system

  24. Venera landers Venera-13 landing sequence (By source) Venera 13 lander & orbiter(By source) (By source) Surface of Venus by Venera-13 (By source)

  25. Airbags Luna-9/13 landers cushioning with airbags. (By source) MER landing sequence(By source) MER airbags(By source)