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milstar: http://www.ucsusa.org/assets/documents/nwgs/section_6.pdf To maneuver, a satellite in orbit must use rocket engines (thrusters) to change the magnitude or direction of its velocity. Because the orbital speed of satellites is so large, the velocity changes required for maneuvering may also be large, requiring the thrusters to use large amounts of propellant. How much and how quickly a satellite can maneuver depends on the amount and type of propellant it carries. There are practical limits to the amount of propellant a satellite can carry since it increases the total mass that must be launched into orbit. These constraints on maneuvering in space have important consequences for satellite operations. This section discusses the different types of satellite maneuvers and the changes in satellite velocity required for each. Section 7 outlines the amount of propellant required for these maneuvers. Three basic maneuvers are used to change orbits: (1) changing the shape or size of an orbit within the orbital plane; (2) changing the orbital plane by changing the inclination of the orbit; and (3) changing the orbital plane by rotating the plane around the Earth’s axis at constant inclination. (Recall that all satellite orbits lie in a plane that passes through the center of the Earth.) We discuss each of these in more detail below, as well as several common orbital changes that use these basic maneuvers. Maneuvers within the orbital plane allow the user to change the altitude of a satellite in a circular orbit, change the shape of the orbit, change the orbital period, change the relative location of two satellites in the same orbit, and de-orbit a satellite to allow it to return to Earth. A velocity change is typically referred to as delta-V, or DV, since the term “delta” is commonly used in technical discussions to indicate a change in some quantity. In addition, as Section 7 shows, generating a velocity change of 2 km/s ################################################ with conventional propulsion technologies would require a satellite to carry its own mass in propellant—thus doubling the mass of the satellite. ############################################# Table 6.1. This table shows the change in satellite velocity (DV) required for various types of maneuvers and activities in space, where Dq is the change in inclination. Type of Satellite Maneuver Required DV (km/s) Changing orbital altitude within LEO (from 400 to 1,000 km) 0.3 Stationkeeping in GEO over 10 years 0.5–1 De-orbiting from LEO to Earth 0.5–2 Changing inclination of orbital plane in GEO by Dq = 30° 2 by Dq = 90° 4 Changing orbital altitude from LEO to GEO (from 400 to 36,000 km) 4 Changing inclination of orbital plane in LEO by Dq = 30° 4 by Dq = 90° 11 These numbers are calculated in the Appendix to Section 6. (LEO = low earth orbit, GEO = geosynchronous orbit)
milstar: fig 7.5 124/142 Skorost i skorost w apogee to ranga 1000 km - w apogee 600-700 m/sek , na zemle 3000 m/sek 2000 km - 1100 m/sek, 4000 m/sek 3000 km -1650 m/sek ,4800 m/sek 4000 km - 2100 m/sek ,5300 m/sek 5000 km - 2600 m/sek ,6000 m/sek 6000 km - 3000 m/sek ,6200 m/sek 7000 km- 3500 m/sek ,6600 m/sek 8000 km -4000 m/sek ,6800 m/sek 9000 km - 4200 m/sek ,7000 m/sek 10000 km -4700 m/sek ,7200 m/sek 11000 km - 5000 m/sek ,7300 m/sek 12000 km -5300 m/sek ,7500 m/sek 15000 km - 6200 m/sek ,7800 m/sek 19000 km -7700 m/sek ,7900 m/sek
milstar: fig 7.6 1257143 apogej ot dalnosti 1000 km -230 km 2000 km -430 km 3000 km -640 km 4000 km -800 km 5000km -960 km 6000 km -1100 km 7000 km -1200 km 8000 km -1260 km 9000 km -1300 km 10000 km -1320 km 11000 km -1300 km 12000 km -1260 km 15000 km -960 km 18000 km -430 km 19000 km -230 km
milstar: Fig 7.7 wremja to dalnosti 1000 km -500 sek 2000 km -700 sek 3000 km -880 sek 3500 km -1000 sek 5000 km -1260 sek 8000 km -1700 sek 10000 km -1900 sek 19000 km -2500 sek
milstar: Fig 7.16 /154 dlja 7000 m/sek ,25° ygol wxoda s dalnostju 8910 km http://www.scribd.com/doc/53416290/Atmospheric-Re-Entry-Vehicle-Mechanics Effect wraschenija Zemli na schirote 45 ° dalnost " - +" 550 km w zawisimosti ot azimuta -90° ... +90° Azimut = 0° pri naprawleniii ot tochki starta na NORD Xoroscho izwestno chto pri zapuske sputnika s Ekwatora wostochnoe naprwalenie ( sowp .s napr. wraschenii Zemli) ############################################################################ imeet preimuschestwa ################## Dalnost ballisticheskix raekt sootwestwenno boslche pri zapuske w wostochnom naprawlenii ############################################################ Dlaj toj ze samoj otnostitelnoj skorosti ,potr. energii wiigrisch ot yvelcihenija absoljutnoj skorosti i ymenschenija flight path angle (ygol soda ) Dannij effekt bolee znachitelen na ekwatore Otnositelno zwezd absoljutnij rang yvelichiwaetsja ----------------------------------------------------------- Xotja za wremja poleta fiks. punkt na Zemle smeschaetsja w wostochnom naprawlenii ,absoljutnaja dalnost wische ... Resultat -ywelichenie dalnosti mezdu 2-mja fiksirowannami punktami Dlja sputnikow danni jeffekt esche wische ... Na ekwatore -1200 km + 1400 km
milstar: Ballistic missile trajectory prediction using a state transition matrix http://faculty.kfupm.edu.sa/AE/aymanma/images/Ballistic%20missile%20trajectory%20prediction%20using%20a%20state%20transition%20matrix.pdf A method for the determination of the trajectory of a ballistic missile over a rotating, spherical Earth given only the launch position and impact point has been developed. The iterative solution presented uses a state transition matrix to correct the initial conditions of the ballistic missile state vector based upon deviations from a desired set of final conditions. A six-degree-of-freedom simulation of a ballistic missile is developed to calculate the resulting trajectory. Given the initial state vector of the ballistic missile, the trajectory is simulated and the state transition matrix propagated along the trajec- tory to the impact point. The error in the final state vector is calculated and elements of the initial state vector are corrected using the state transition matrix. The process is repeated until the ballistic missile impacts the target location within a pre- defined miss distance tolerance. The result of this research is an analysis tool which accurately solves for the initial state vector of the ballistic missile.
milstar: Na wisotax nize 50 km -lobowoe soprotivlenie (drag) stanowitsja wische chem wes boegolowki str.149/168
milstar: http://www.scribd.com/doc/53416290/Atmospheric-Re-Entry-Vehicle-Mechanics 168/179 table 9.2 dlja betta 10000 kg/kw.metr i nosa 25 mm (radiusa ?) skorost kasanija s zemlej 10000 km - 1916 metr/sec 8000 km -2211 metr/sec 6000 km -2347 metr /sec = M6.9 4000 km - 2273 metr/sec 9.3. skorost dlaj 6000 km dalnosti w zawisimosti ot ballisticheskogo koef . beta 20000 -11.2 M 3818 metr/sec 10000 -6.9 M 2347 metr/sek 5000 - 2.7 926 metr /sek 9.4 wremja reentry wisota 120 km do kasaniya Zemli 2000 km -55.8 sek 4000 km - 43.5 sek 6000 km -42 sek 8000 km -44.6 sek 10000 km -50.7 sek 12000 km -62.3 sek S tochki zrneija perexwata ( minimlno zatr. traektorija i pri prochix rawnxi) naibolee slozni boegolowki s wisokim Beta =20000 na distanzii 6000 km ( ili blizkix 4000 km i 8000 km) , kotorie imejut max skorost kasanija s Zemlej i minimalnoe wremja reentry s wisoti 120 km =42 sek pri beta 20000 na dist 6000 km - 3818 metra/sek pri beta 10000 na dist 6000 km -2347 metra/sek Minimalno zatratno -start w wotochnom naprawlenii na odnoj shirote s celju ################################################## S rajonow Saxalina ,Kamchatki po celjam na zapadnom poberez'e USA
milstar: THE DESCRIPTIVE GEOMETRY OF NOSE CONES http://www.if.sc.usp.br/~projetosulfos/artigos/NoseCone_EQN2.PDF Nose cone design http://en.wikipedia.org/wiki/Nose_cone_design
milstar: http://www.scribd.com/doc/73116189/Ballistic-Glide-Re-Entry-Vehicle-BGRV-and-Indian-Missile-Program Page:905 Feb, 2011 BGRV •Aerodynamics –Long, high aspect ratio body –Slight angle of attack to generate lift. –Control surfaces: fins | paddles | small gas thrusters •Flight –On board flight control and navigation: Hi Accuracy –~ 25 G maneuvering –Porpoise maneuver in pitch axis –“S” shape maneuver in azimuth plane –Rolling to spread heat and reduce thermal shielding
milstar: Srawnenie konstrukzij boewix blokow i traektorij MARV str. 21-27 ########################################### 1. Pershing-2 2.Agni BGRV 4.2 metra dlinnoj ruli ,dwigatel ,yabch 3.Kitaj/Pakistan -M11 wes yabch okolo 35 % ot 1762 pounds 4. Topol-M http://www.scribd.com/doc/73116189/Ballistic-Glide-Re-Entry-Vehicle-BGRV-and-Indian-Missile-Program
milstar: The maneuverable reentry vehicle (abbreviated MARV or MaRV) is a type of ballistic missile warhead capable of shifting targets in flight. Refer to atmospheric(? to awtora postinga ) reentry. ----------------------------------------------------------- There are several types, of which examples include: 1. Agni-V -bolee 5000 km 2. Trident -the version designed for the Trident missile, which had to be able to evade Soviet anti-ballistic missile systems.(Evader Mk.500) 3. MGM-31C Pershing II the active radar terminal-guidance version with pinpoint accuracy for the MGM-31C Pershing II missile 4. Pakistan -Shaheen IA 5. Pakistan -Shaheen-II 2500 km 6. Kitaj B-611 -400 + km 7.DF-15 8. the high hypersonic land-based anti-ship ballistic missile variant of the DF-21 2500 -2700 km 9.DF-31 11000 -12000 km,42 tonni 3 MaRV 20 kt,90 kt,150 kt 10DF-41 12000 km ,30 tonn 11. JL-2 do 14000 km s ymensch. nagruzkoj ,42 tonni 12.R-29RMU2 Liner 13. RSM-56 Bulava the warheads used by the Topol-M and RS-24 missile which are designed to defeat US ABM systems. (awtor postinga SS-27 Sicle B -1 Marv i SS-x-29 3 MaRv) http://en.wikipedia.org/wiki/B-611
milstar: Mk. 500 Evader Маневрирующая боеголовка Старт программам разработки маневрирующих боевых блоков был положен в конце 60х годов. В отличие от обычных боеголовок завершающих полет по баллистической траектории, Evader должен был быть способен совершать предварительно подготовленные маневры на атмосферном участке полета, что позволило бы эффективно противостоять противоракетной обороне противника. Mk. 500 имела простую навигационную систему и представляла из себя конус со скошенной носовой частью. Создаваемая за счет этого подъемная сила контролировалась качанием корпуса за счет передвижения внутренней инертной массы, роль которой выполнял блок электроники. Это было достаточно примитивное устройство при разработке которого главным требованием было способность совершать манер противоракетного уклонения, в то время как точность не являлась критичной. Тем не менее, дальнейшие тесты показали, что потеря точности по сравнению с базовым боевым блоком была вполне приемлемой, и более того, последующие исследования подсказали путь использования маневрирующей боеголовки для существенного улучшения точности, когда достигнув цели боеголовка начинала пикировать вертикально вниз, тем самым устраняя разброс создаваемый разницей в скорости подрыва зарядов. MaVR Evader была одна из наиболее продвинутых тестовых программ проведенных Lockheed совместно с General Electric. Все шесть тестов (6 марта, 9 мая, 22 августа, 10 сентября 1975 года, и 23 января 1976 года) произведенных на военно-воздушной базе Vandenberg использовавших в качестве ускорителей Minuteman и Trident I были признаны успешными. В результате тестов была продемонстрирована возможность реализации маневрирующего боевого блока, были отработаны системы управления и наведения, отработана совместимость с системой Trident I. Несмотря на успешность тестов, Mk. 500 не выпускалась серийно. Предоставляемые возможности высокоточного поражения защищенных целей насторожили членов конгресса, опасавшихся что данная система может послужить предметом эскалации и спровоцирует новый виток гонки вооружений. Разработка была заморожена с возможностью начать серийное производство Mk. 500 в сжатые срока в случае возникновения такой необходимости. Программа AMaRV (Продвинутый Маневрирующий Боевой Блок разработки McDonnell Douglas) Целью опытной программы AMaVR было улучшение качеств заложенных в маневрирующий боевой блок предыдущего поколения (Mk. 500), позволяющих как выполнять маневры уклонения на от более совершенных перехватчиков на большей высоте и скорости, так и повысить точность по сравнению с традиционными баллистическими боеголовками. Проект был рассмотрен в январе 1976 года, и в сентябре того же года был одобрен контракт на строительство четырех блоков и проведение полетных испытаний, состоявшихся 20 декабря 1979, 8 октября 1980 и 4 октября 1981 года с использованием ракеты Minuteman I. Боеголовка имела массу около 470 кг, с носовой частью в виде усеченного конуса с радиусом носовой части 2.34 см и углом между высотой и образующей в 10.4°, радиусом промежуточного сечения 14.6 см, с уменьшением угла до 6°. Полная длина составляла 2.079 метра. Система управлялась посредством четырех отклоняемых закрылков с гидравлическим приводом, использовала инерциальную систему наведения (тестировалась с использованием лазерных гироскопов) и систему астрокоррекции; позволяла использовать новые углы атаки, диапазоны скорости и ускорения. Прецизионный управляемый боевой блок (PGRV Mk. 600) Разработка прецизионного управляемого боевого блока (PGRV) началась в 1976 году. Боевой блок в дополнение к технологиям отработанным в программе AMaRV, использует датчики терминального наведения, позволяющих корректировать траекторию на заключительном этапе полета. Использование предыдущих технологических наработок, позволило создать систему существенно повысившую возможности поражения высокозащищенных целей, или при использовании зарядов малой мощности, достичь замечательных характеристик уничтожения целей при минимальном сопутствующем ущербе. Боеголовка получила наименование Mk. 600 и является альтернативным оснащением для ракет Trident II. При подготовке текста использовались следующие материалы: Nuclear Weapons Databook, Volume I, p. 110 From Polaris to Trident: The Development of US Fleet Ballistic Missile Technology, p. 134 Dynamics of Atmospheric Re-Entry, American Institute of Aeronautics and Astronautics http://www.ssp.navy.mil/about/history_chronology_66-75.shtml http://www.ssp.navy.mil/about/history_chronology_76-85.shtml и другие
milstar: http://www.dtic.mil/cgi-bin/GetTRDoc?AD=ADA315439&Location=U2&doc=GetTRDoc.pdf Pershing II missile. After the warhead reenter the atmosphere, conducts the lift - drag flying maneuver. On one hand, it is advantageous to avoid being intercepted by the air defense, enhance the surprise attack ability, on the other hand, it will slow down the warhead, make the warhead fall to the target almost vertically at the end of the trajectory, create the favorable condition for the scene matching terminal guidance. When the warhead fall down to about 8 Km - 10 Km high, the radar area correlation terminal guidance will start working. At that time, the real aperture image radar antenna will begin a circular image scanning, the scanning frequency is 2 cycles per second, the center is vertical to the horizon, thetargetareaisthedeadcenterunderthewarhead. Thecircularradar image of target area is processed through geometric distortion correction, coordinate transformation and image pre-processing, the scanning frequency is 2 cycles per second, the center is vertical to the horizon, thetargetareaisthedeadcenterunderthewarhead. Thecircularradar image of target area is processed through geometric distortion correction, coordinate transformation and image pre-processing, and then is correlatively compared with the pre-stored reference images in the correlation computer, the best matching point at which the real time radar image coincide in the reference image is found, the precise position of the warhead relating to the target is calculated. This position informationisusedbythecontrolsystemtocorrectthetrajectory. The time for finishing image correlation is about 1 second, in which the former 0.5 second is used by the radar antenna to scan a circle to get the real time radar image, the later 0.5 second is used for correlation processing. Thewholesectionofradarareacorrelationterminalguidance need 3 to 4 times such revision process till the warhead is 900 meter abovetheground. Afterthatthewarheadwillfalldirectlytowardsthe target. These several times of correlation locating and trajectory revision, lowered the Pershing II missile's CEP of targeting to only 30 meters. The radar area correlation terminal guidance system that Pershing II missile uses is developed by GoodYear Corporation. It includes a three degree of freedom stabilized image radar antenna, radar transmitting and receiving system, high speed correlation processor, power supply system, digital reference image, correlation processing software, radar antenna cover which is also acting as warhead's cowl. The main part of the radar area correlation equipment weights about 57 Kg without the radar antenna cover. This part is placed in the front of the warhead, among which the three degree of freedom antenna is in the antenna coverwhichhasthegoodperformanceofwavetransmitting. Theimage radar antenna is in the shape of cutting paraboloid, adopts the deviation focus forward feed horn illuminator to form a 2.2 * 22 degree sector wave
milstar: and assure that pitching wave beam (22 degree direction) deviates from the antenna rotation scanning center in a fixed angle. When the antenna scans the target area to get image, the rotation center of the antenna need to track the vertical line because of the different warhead postures. It means that the antenna need to do position and pitching follow-up. Thus this radar image antenna is a kind of stable antenna equipment which has three degree of freedom of position, pitching and rotation and equipped with three follow-up control systems. The image radar that Pershing II warhead terminal guidance system uses is a real apertureradarwhoseworkingfrequencyisinJband. Thesenderinthis radar is a incoherent pulse magnetron sender, the success rate of sending pulse peak value is about 60 km. In order to improve the performance of this radar, the frequency sudden changing technology is used. The received ground echo wave signal is processed by direct frequency mixing, local oscillator uses sudden frequency changing tracking local oscillator with appropriate frequency. The logarithm intermediate frequency amplifier preserves the reflection system information of different land objects in big area. Besides the pre-processing and output data processing, the main computing task of the correlation computer is correlation operation. For example, the pixels of the real time radar image are 128 * 128 matrix, the pixels of the reference image are 256 * 256 matrix, the computing quantity of finishing a whole image searching need 270 million operations, if it is necessary to finish this amount of computing within 0.5 second, the computer speed will be more than 500 millionoperationspersecond. Eventhesimplestcorrelationalgorithmis used, such as Mean Absolute Difference (MAD) algorithm, and with effective speedy searching algorithm which cut the computing burden significantly, the computer's speed should be more than 50 million operations per second. In the 70's, because of the application difficulty of the computer technology in the missile, an optical correlation device was used instead. Later, the all digitized correlation processor took place of the optical correlation device. To achieve the processing speed of 50 million operations (addition) per second in the missile, it is impossible to imagined with a single CPU computer at that time. So, we can deduce a conclusion that its correlation computer uses multi CPU parallel connectingtechnology,orarraycomputertechnology. Theradarreference images that this radar area correlation processor uses are made by the automatic reference map generating equipment which is developed by the GAC Corporation. ThisequipmentusesthestandardizeddatabaseDLMSproduced by National Defense Mapping Bureau of the US, the designated target area, and the radar parameters in the missile to generate the radar reference image which can be stored onto the memory in the missile's correlation computer.
milstar: http://www.nap.edu/openbook.php?record_id=12061&page=206 The Why and How of Boost-Glide Systems Given the prominence of the boost-glide technology in some of the options under consideration in this report, it is useful to include an appendix explaining semiquantitatively what the technology can and cannot accomplish, its relation to the fractional orbit bombardment systems (FOBSs) technology discussed during the 1960s and 1970s, and some of the technical challenges involved. Another issue is the extent to which such vehicles can be expected to defeat “garden-variety” and advanced air defenses. A boost-glide vehicle (BGV), or “lifting body” without propulsion, can be used to extend the range of a ballistic-missile payload beyond the purely ballistic range. It can also be used for out-of-plane or “dogleg” maneuvers to avoid over-flight of certain areas or to allow the dropping of initial rocket stages into the sea or into another body of water not under the ballistic path. The space shuttle on reentry is an example of a hypersonic lifting body. First consider the BGV for relatively short-range systems—up to a few thousand kilometers in range—in the approximation of a flat Earth. An important simplification arises from the fact that the atmosphere is shallow; the air density falls by a factor e = 2.72 for each 8 km increase in altitude. As is the case with a normal glider, the aerodynamic support of the vehicle against gravity (the “lift” L) is accompanied by “drag due to lift,” as characterized by the lift-to-drag ratio (L/D); for a clean subsonic glider aircraft this may be as much as 40, but for a hypersonic lifting body an L/D = 2.2 is an achievement. In the numerical examples, it is assumed that L/D = 2.2. With the glider aircraft or for a powered vehicle that has run out of fuel and that is gliding for as long a distance as possible, the drag, D, extracts energy from the vehicle—from its store of kinetic energy MV2/2 and potential energy MgH;
milstar: This ballistic reentry wastes the kinetic energy of the RV at the time of reentry, whose velocity is the minimum required to achieve the desired range in the first place. This is the best that could be done if Earth had no atmosphere. But it does, and in principle the RV could be designed as a lifting body for the hypersonic regime, and if the thermal insult could be managed it could transition in the upper regions of the atmosphere to near-horizontal flight, and then use lift and change of altitude, air density, and change of angle of attack to support the RV weight for a substantial range extension beyond the purely ballistic trajectory. This approach was validated decades ago by flights of the Mk-500 “Evader” RV. Successful implementation of boost-glide technology could yield additional benefits for the prompt global strike mission by means of the ability to maneuver and thus to aid in avoiding undesired overflight of various countries. The launch would be similar to that for a minimum-energy trajectory—that is, maximum range for a given missile—typically with a high apogee and the transition on ballistic reentry to either level or phugoid (porpoise-like) flight—in which the RV bounces in and out of the atmosphere several times and supports its weight by aerodynamic lift only a relatively small fraction of the time, say 10 percent. Supporters of the BGV often argue that this phugoid flight provides range extension at little cost, because for much of this flight—between bounces—the drag is almost zero. It is important to recognize that there is “no free lunch” in phugoid flight, ############################################### (str.207) because the lift averaged over this portion of the flight is precisely the weight of the vehicle, and so the time-average lift (and drag) are the same as if the RV were flying at steady altitude and speed in order to maintain the same average aerody namic lift. The average lift must be equal to the weight of the vehicle: W = gM; the average drag is thus the weight divided by (L/D). On the assumption of constant L/D, it turns out that there are simple closed-form formulas not only for the glide portion of flight but also for the velocity and kinetic-energy loss in the transition from ballistic flight to glide.
milstar: http://www.nap.edu/openbook.php?record_id=12061&page=210#p200161e09960210001 FIGURE G-1 Reentry trajectories for L/D = 2.2. Note that the last 1,000 km or more of the reentry trajectories are identical for Earth and flat Earth. SOURCE: Data for initial conditions (7 km/s, −10° grazing) provided by G. Candler, University of Minnesota, personal communication to the committee, September 17, 2007. As indicated in Figure G-1, much of the range benefit from boost-glide in general and phugoid flight in particular is only available on the round Earth and with near-orbital initial speed of the RV. For speeds in the upper atmosphere comparable to the orbital velocity in low Earth orbit (LEO), almost no aerodynamic lift is necessary, so the glide range can be astonishing—say, on the order of 13,000 nautical miles (nmi) in some cases. Since Earth’s circumference is 360 degrees of arc and each arc minute is 1 nmi, the circumference of the world is 21,600 nmi (precisely 40,000 km, by the definition of the meter).
milstar: Some of the proposals for long-range boost-glide vehicles enter the glide phase at angles from the horizontal two to four times smaller than the 10° example used here, and at speeds considerably closer to orbital speed of 7.90 km/s (25,920 ft/s) than the example of 7 km/s used here. They are essentially “fractional orbital bombardment vehicles” ####################################### str 214 http://www.nap.edu/openbook.php?record_id=12061&page=214 with essentially infinite “range extension” and substantial cross-range maneuver capability.
milstar: BGVs of longest range start at essentially orbital speed of 7.9 km/s and thus have more kinetic energy to dissipate than do ICBM RVs. The RV, however, traverses the 8 km “scale height” of the atmosphere at an angle to the horizontal of 22°, in a few seconds, while the BGV supports itself aerodynamically for 10,000 km at near-orbital speed for 1,200 s. The heating due to lift is concentrated on the lower surface of the BGV rather than uniformly around the axis of the RV, usually resulting in a very thick layer of ablative material on the lower surface of the BGV. The function of this inner layer is simple insulation rather than ablation, and so the thermal protection Indeed, much of the protection system could be in the form of non-ablating material such as the “tiles” on the space shuttle. The intense heating of the BGV during the whole of the glide phase provides a strong infrared signal to defensive systems ########################################################################## equipped to detect it or to use it for an infrared homing intercept. ################################ A simple terminal maneuver for a ballistic missile will allow it to deny sanctuary to structures and locations shielded by a near-vertical bluff. At intermediate range this can require a 45° maneuver that with an L/D = 2.2 would (according the example following Equation G-2) result in a reduction of warhead speed to 0.6998 of the initial speed. If performed at 10 g transverse acceleration (0.098 km/s2), the maneuver could take on the order of 30 s; an alternative would be to have a high-drag RV to greatly reduce speed to, say, Mach 3 (1 km/s), so that a 45° maneuver could be accomplished in a few seconds (slowdown to turn). The simple kinematic considerations of this appendix indicate the value of the engineering design of a variable-geometry RV, and the competition between the longer-term “better” and the earlier and perhaps “good enough.”
milstar: http://www.ausairpower.net/APA-Sov-FOBS-Program.html The Soviet Fractional Orbital Bombardment System Program The Fractional Orbital Bombardment System (FOBS) as it was known in the West, was a Soviet innovation intended to exploit the limitations of US BMEW radar coverage. The idea behind FOBS was that a large thermonuclear warhead could be inserted into a steeply inclined low altitude polar orbit, such that it would approach the CONUS from any direction, but primarily from the southern hemisphere, and following a programmed braking manoeuvre, re-enter from a direction which was not covered by US BMEW radars. The first warning the US would have of such a strike in progress would be the EMP transients produced by the nuclear devices initiating over their programmed targets in the CONUS. The missile’s flight profile comprised four phases – boost phase, orbital phase, braking phase and finally, the re-entry phase. The weapon’s 1,700 kg orbital stage was designated the 8F021 OGCh, which comprised a fuselage, an instrument section with an inertial guidance system, the de-orbit engine section, and an 8F673 ~5 Megatonne nuclear warhead section. The 8F021 would, as it neared the de-orbit manoeuvre entry point, start the AT/UDMH liquid fuelled de-orbit engine turbopump using a solid propellant gas generator. Exhaust gasses from the turbine were used for vehicle attitude control, using a 4 + 4 thruster arrangement. This de-orbit engine design later formed the basis of the Tsiklon 3 ELV S5.23/RD-861 third stage orbital engine, rated at 78.710 kN / 17,695 lbf. The cited CEP for the RV was 1.1 km. Conceived at the peak of the Cold War, the Soviet FOBS effort showed the extreme lengths to which the Soviets were prepared to go in order to gain a decisive advantage over the West in a nuclear confrontation. The usefulness of the FOBS declined very rapidly, as the US deployed early warning satellites capable of tracking missile launch signatures, and the expanded coverage BMEWS network, with the new phased array AN/FPS-115 PAVE PAWS detection and precision tracking radars.
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