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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 7.16 /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: Compared with forward-deployed systems, prompt strike from within the CONUS requires a larger rocket and results in a flatter reentry angle and higher reentry speeds. ####################################### The larger systems allow larger payloads than those that could be carried by CTM or SLGSM; however, the higher reentry speed renders more difficult the control, guidance, and navigation needed for accurate targeting, and the long exposure to high heat flux complicates thermal management. Nevertheless, developing a longer-range glide capability for the reentry vehicle in the form of a 161-inch boost-glide vehicle (larger than the modification proposed for the SLGSM) is the basis for the land-based options listed in Table 4-1. Additional details on the proposed boost-glide systems are provided below. The reader is also referred to Appendix G for additional details on the characteristics of the boost-glide trajectory. http://www.nap.edu/openbook.php?record_id=12061&page=113
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 65 ° dalnost " - +" 550 km w zawisimosti ot azimuta -90° ... +90° ################################################################### na schirote 45° +970 km - 840 km ########################## 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 dannij effekt esche wische ...
milstar: http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19650069911_1965069911.pdf
milstar: DEPRESSED TRAJECTORIES • Latedetectionandrelativelylowinterception altitudes (attacker’s advantage) • RV appears without its decoys (defense advantage) • Largerdispersionofimpactpoints (as compared with Minimum Energy trajectories) • RVneedsspecialdesign (aerodynamic heating problems) • Prediction of PIP less accurate (as compared with Minimum Energy trajectories) LOFTED AND DEPRESSED TRAJECTORIES For each range, there are infinite combinations of burnout velocity and burnout trajectory angle 1) Burnout velocity is known (i.e. rocket motor is given) (2) A range (less than minimum energy range) has been specified There are two trajectories leading to this range: (a) Lofted trajectory (b) Depressed trajectory
milstar: http://www.princeton.edu/sgs/publications/sgs/pdf/3_1-2gronlund.pdf The most energy-efficient trajectory for flying a ballistic missile over a given range carries it high above the atmosphere; the maximum range for a given missile is achieved by traveling on such a trajectory. If a missile is flown over shorter ranges, the excess energy can be used to fly less energy-efficient trajectories, such as low-apogee or "depressed" trajectories. Missiles flown on a depressed trajectory (DT) can have significantly shorter flight paths, and therefore significantly shorter flight times, than those flown on a standard trajectory of the same range. This is especially relevant for sea-launched ballistic missiles (SLBMs) which could, in principle, be brought close to the territories of the US or of the Commonwealth of Independent States (CIS). We will see that a Trident II SLBM on a depressed trajectory could travel 1,850 kilometers in roughly 7 minutes (rather than the 12.5 minute flight time on a standard trajectory), and could therefore reduce the survivability of a large fraction of the bomber forces that would not be vulnerable to attack by SLBMs on standard trajectories. High accuracy is not required for attacking bombers as they are disabled by overpressures of 2-5 psi (14-34 kPa),10 which can be produced at relatively large distances from a nuclear blast. Warheads of 100 to 500 kilotons can produce these overpressures at distances of 2 to 5 kilometers from bombers flying at altitudes of several kilometers, and at roughly twice these distances for bombers on the ground. 11 We will see below that although accuracies of ################################ SLBMs are degraded by flying them on depressed trajectories, ############################################ they would be adequate to attack such soft targets. 12 The development of maneuvering RVs (MaRVs) for DT SLBMs COt lead to a combination of short flight times and sufficient accuracy to atta hardened targets, thereby threatening even hardened command posts ve early in a conflict and silo-based ICBMs. Mobile ICBMs are relatively soft tl gets that rely on dispersal for their survival; if deployed in garrison th would be vulnerable to attack by DT SLBM
milstar: There are four main questions to be addressed in considering trajectories wi1 low apogees. 1.First, because a DT SLBM would spend a longer time in tl denser parts of the atmosphere than an SLBM on a standard trajectory, wou: this lead to increased aerodynamic stress or heating on the booster, and woul a new booster be required? 2.Second, would the additional time spent in tt. atmosphere lead to increased heating of the RV, thereby requiring a new R design? 3.Third, what flight times are achievable? 4.Finally, to what extent wou] the accuracy of an SLBM on a depressed trajectory be degraded, and wou] the development of a precision-guided RV (PGRV) be required? If the develoJ ment of a new booster or RV is required, constraining DT capabilities via arm control would be relatively straightforward. If no new development is needec acquiring DT capabilities would probably entail only flight testing of existin boosters and RVs on the new trajecto
milstar: For each flight range and altitude at burnout, there is a unique ballistic missile trajectory that is most energy efficient. This is known as the minimum- energy trajectory (MET). If flown over less than maximum range, an SLBM can use its excess fuel to fly on a less energy-efficient, lower- or higherapogee trajectory. In routine operations, SLBM trajectories may be lofted or depressed slightly with respect to the minimum-energy trajectory for a variety of reasons. However, "depressed trajectory" refers to trajectories that are depressed substantially below the MET: these trajectories have reentry angles. of roughly 5-100 for a 1,850 kilometer range, compared to roughly 400 for a MET of comparable range, or roughly 300 for a 7,400 kilometer MET. Figure 2 shows a minimum-energy and two depressed trajectories with ranges of 1,850 kilometers.
milstar: http://www.princeton.edu/sgs/publications/sgs/pdf/8_1Zarchan.pdf
milstar: Trident -2 Mk-4 14*100 kt ,7400 km ,29,2 minut ,apogej -1340 km ,burnout ckosrost - 6.3 km/sek Ballistic coefficient 8600 N/m2 Weight of RV 90 kilograms Weight of bus 1,100 kilograms Length -1.3 metra Nose radius -0.045 metr Base radius - 0.2 metr Half angle -7 degress http://www.princeton.edu/sgs/publications/sgs/pdf/3_1-2gronlund.pdf Trident -2 Mk-5 ,8*475 km ,7400 km Ballistic coefficient 12000 N/m2 Weight of RV 180 kilograms Weight of bus 1,100 kilograms Length -1.5 metra Nose radius -0.04 metr Base radius - 0.26 metr Half angle -8.5 degress
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: для баллистического коэффициента 10000 килограмм/кв.метр и носа радиусом 25 миллиметров скорость касания с Землей ---- 10000 км - 1916 метр в сек 8000 км -2211 метр в сек 6000 км -2347 метр в сек= M6.9 4000 км - 2273 метр в сек скорость касания с Землей в зависимости от баллистического коэффициента для дальности 6000 километров --------- beta килограмм/кв.метр 20000 -11.2 M 3818 метр в сек 10000 -6.9 M 2347 метр в сек 5000 - 2.7 M 926 метр в сек время входа с высоты 120 километров до касания с землей ------------------------------------- 2000 км - -55.8 сек 4000 км -- 43.5 сек 6000 км -42 сек 8000 км --44.6 сек 10000 км -50.7 сек 12000 км -62.3 сек
milstar: На ракетах Р-29РМ (РМУ) впервые наряду с астрокоррекцией применена радиокоррекция по навигационным спутникам Земли. Обеспечена стрельба из высоких широт Арктики и по настильным траекториям с малым подлетным временем. --------------------------------------- Трехступенчатая схема ракеты не имеет аналогов среди жидкостных боевых ракет как у нас, так и за рубежом. Ракеты обладают модернизационным потенциалом, реализация которого благоприятно сказалась на поддержании боевых свойств морских стратегических ядерных сил за счет установки более эффективных боевых нагрузок, в том числе средств противодействия противоракетной обороне, в последующие годы. Владимир Григорьевич Дегтярь - доктор технических наук, член-корреспондент РАН, академик РАРАН, генеральный директор, генеральный конструктор ОАО "ГРЦ Макеева"; Рэм Никифорович Канин - кандидат технических наук, ведущий научный сотрудник ОАО "ГРЦ Макеева". http://nvo.ng.ru/armament/2012-11-02/1_rockets.html
milstar: http://www.ucsusa.org/assets/documents/nwgs/space_weapons_section_7.pdf Delta V 2km/sek 1 tonna sputnik +0.9 tonni toplivo 1 km/sek =0.4 1km/sek 1 tonna 0.4 tonni For example, to carry out a maneuver requiring a ∆V of 2 km/s, the propellant mass Mp required for this maneuver is 0.9 times that of the satellite itself, that is, the propellant nearly doubles the total mass that must be placed in orbit. In other words, a satellite with a mass of one ton (excluding the propellant for this maneuver) would need to carry 0.9 tons of propellant to provide the ∆V for this maneuver. -------- Changing orbital altitude within LEO (from 400 to 1,000 km) delta V = 0.3 km/s 0.1 Mp/Ms De-orbiting from LEO to Earth 0.5–2 0.2–1 http://www.ucsusa.org/assets/documents/nwgs/space_weapons_section_7.pdf
milstar: Недостаток неядерных решений - 1...резко падают баллистические коэффициенты и соответственно скорость 2...растет ЭПР боевого блока для баллистического коэффициента 10000 килограмм/кв.метр и носа радиусом 25 миллиметров скорость касания с Землей ---- 10000 км - 1916 метр в сек 8000 км -2211 метр в сек 6000 км -2347 метр в сек= M6.9 4000 км - 2273 метр в сек скорость касания с Землей в зависимости от баллистического коэффициента для дальности 6000 километров --------- beta килограмм/кв.метр 20000 -11.2 M 3818 метр в сек 10000 -6.9 M 2347 метр в сек 5000 - 2.7 M 926 метр в сек время входа с высоты 120 километров до касания с землей ------------------------------------- 2000 км - -55.8 сек 4000 км -- 43.5 сек 6000 км -42 сек 8000 км --44.6 сек 10000 км -50.7 сек 12000 км -62.3 сек
milstar: http://www.fas.org/rlg/030522-space.pdf
milstar: Representative Reentry Vehicle Characteristics Nose radius (cm) 1.98 Base radius (cm) 22 Length (cm) 152 Mass (kg) 92 Drag coefficient 0.1 Ballistic coefficient (β)-60000 SOURCE: Regan (1984), p. 333. tungsten http://www.rand.org/content/dam/rand/pubs/monograph_reports/2011/RAND_MR1209.pdf
milstar: http://www.academypublisher.com/proc/iscsct09/papers/iscsct09p448.pdf
milstar: http://www.chinasecurity.us/pdfs/others/Hagt&Durnin.pdf
milstar: http://www.scribd.com/doc/53416290/Atmospheric-Re-Entry-Vehicle-Mechanics Minimal energy trajectory ################ str.124/142 Fig. 7.4 Initial flight path angle as function of range s zemli 2000 km -40° 4000 km -36 ° 6000 km -32 ° 8000 km -27 ° 9000 km -25° 10000 km -22.55° 11000 km -20 ° 18000 km -5 ° ######### 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 ###################### 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 ########### 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
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