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76-508: The Transfer Orbit Stage ( TOS ) was an upper stage developed by Martin Marietta for Orbital Sciences Corporation during the late 1980s and early 1990s. The TOS was designed to be a lower-cost alternative to Inertial Upper Stage and Centaur G upper stages. The TOS was designed to be deployed by the Titan 34D , Commercial Titan III and Space Shuttle . The main propulsion system of

152-570: A domino effect known as Kessler syndrome . NASA's Orbital Debris Program tracks over 25,000 objects larger than 10 cm diameter in LEO, while the estimated number between 1 and 10 cm is 500,000, and the number of particles bigger than 1 mm exceeds 100 million. The particles travel at speeds up to 7.8 km/s (28,000 km/h; 17,500 mph), so even a small impact can severely damage a spacecraft. [REDACTED]  This article incorporates public domain material from websites or documents of

228-404: A semi-major axis of 8,413 km (5,228 mi). For circular orbits, this in turn corresponds to an altitude of 2,042 km (1,269 mi) above the mean radius of Earth, which is consistent with some of the upper altitude limits in some LEO definitions. The LEO region is defined by some sources as a region in space that LEO orbits occupy. Some highly elliptical orbits may pass through

304-457: A collision risk to the many LEO satellites. No human spaceflights other than the lunar missions of the Apollo program (1968-1972) and the 2024 Polaris Dawn have taken place beyond LEO. All space stations to date have operated geocentric within LEO. A wide variety of sources define LEO in terms of altitude . The altitude of an object in an elliptic orbit can vary significantly along

380-417: A crane. This is generally not practical for larger space vehicles, which are assembled off the pad and moved into place on the launch site by various methods. NASA's Apollo / Saturn V crewed Moon landing vehicle, and Space Shuttle , were assembled vertically onto mobile launcher platforms with attached launch umbilical towers, in a Vehicle Assembly Building , and then a special crawler-transporter moved

456-510: A dragon's head with an open mouth. The British scientist and historian Joseph Needham points out that the written material and depicted illustration of this rocket come from the oldest stratum of the Huolongjing , which can be dated roughly 1300–1350 AD (from the book's part 1, chapter 3, page 23). Another example of an early multistaged rocket is the Juhwa (走火) of Korean development. It

532-408: A higher cost for deployment. Hot-staging is a type of rocket staging in which the next stage fires its engines before separation instead of after. During hot-staging, the earlier stage throttles down its engines. Hot-staging may reduce the complexity of stage separation, and gives a small extra payload capacity to the booster. It also eliminates the need for ullage motors , as the acceleration from

608-415: A higher specific impulse means a more efficient rocket engine, capable of burning for longer periods of time. In terms of staging, the initial rocket stages usually have a lower specific impulse rating, trading efficiency for superior thrust in order to quickly push the rocket into higher altitudes. Later stages of the rocket usually have a higher specific impulse rating because the vehicle is further outside

684-441: A mission is the burn time, which is the amount of time the rocket engine will last before it has exhausted all of its propellant. For most non-final stages, thrust and specific impulse can be assumed constant, which allows the equation for burn time to be written as: Where m 0 {\displaystyle m_{\mathrm {0} }} and m f {\displaystyle m_{\mathrm {f} }} are

760-472: A multistage rocket introduces additional risk into the success of the launch mission. Reducing the number of separation events results in a reduction in complexity . Separation events occur when stages or strap-on boosters separate after use, when the payload fairing separates prior to orbital insertion, or when used, a launch escape system which separates after the early phase of a launch. Pyrotechnic fasteners , or in some cases pneumatic systems like on

836-466: A rocket system will be when performing optimizations and comparing varying configurations for a mission. For initial sizing, the rocket equations can be used to derive the amount of propellant needed for the rocket based on the specific impulse of the engine and the total impulse required in N·s. The equation is: where g is the gravity constant of Earth. This also enables the volume of storage required for

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912-543: A satellite into a LEO, and a satellite there needs less powerful amplifiers for successful transmission, LEO is used for many communication applications, such as the Iridium phone system . Some communication satellites use much higher geostationary orbits and move at the same angular velocity as the Earth as to appear stationary above one location on the planet. Unlike geosynchronous satellites , satellites in low orbit have

988-402: A single rocket stage. The multistage rocket overcomes this limit by splitting the delta-v into fractions. As each lower stage drops off and the succeeding stage fires, the rest of the rocket is still traveling near the burnout speed. Each lower stage's dry mass includes the propellant in the upper stages, and each succeeding upper stage has reduced its dry mass by discarding the useless dry mass of

1064-441: A small field of view and can only observe and communicate with a fraction of the Earth at a given time. This means that a large network (or constellation ) of satellites is required to provide continuous coverage. Satellites at lower altitudes of orbit are in the atmosphere and suffer from rapid orbital decay , requiring either periodic re-boosting to maintain stable orbits, or the launching of replacements for those that re-enter

1140-549: A stable low Earth orbit is about 7.8 km/s (4.8 mi/s), which translates to 28,000 km/h (17,000 mph). However, this depends on the exact altitude of the orbit. Calculated for a circular orbit of 200 km (120 mi) the orbital velocity is 7.79 km/s (4.84 mi/s), but for a higher 1,500 km (930 mi) orbit the velocity is reduced to 7.12 km/s (4.42 mi/s). The launch vehicle's delta-v needed to achieve low Earth orbit starts around 9.4 km/s (5.8 mi/s). The pull of gravity in LEO

1216-433: A subset of LEO. These orbits, with low orbital inclination , allow rapid revisit times over low-latitude locations on Earth. Prograde equatorial LEOs also have lower delta-v launch requirements because they take advantage of the Earth's rotation. Other useful LEO orbits including polar orbits and Sun-synchronous orbits have a higher inclinations to the equator and provide coverage for higher latitudes on Earth. Some of

1292-399: A technical algorithm that generates an analytical solution that can be implemented by a program, or simple trial and error. For the trial and error approach, it is best to begin with the final stage, calculating the initial mass which becomes the payload for the previous stage. From there it is easy to progress all the way down to the initial stage in the same manner, sizing all the stages of

1368-431: Is a commonly used rocket system to attain Earth orbit. The spacecraft uses three distinct stages to provide propulsion consecutively in order to achieve orbital velocity. It is intermediate between a four-stage-to-orbit launcher and a two-stage-to-orbit launcher. Other designs (in fact, most modern medium- to heavy-lift designs) do not have all three stages inline on the main stack, instead having strap-on boosters for

1444-415: Is a safe and reasonable assumption to say that 91 to 94 percent of the total mass is fuel. It is also important to note there is a small percentage of "residual" propellant that will be left stuck and unusable inside the tank, and should also be taken into consideration when determining amount of fuel for the rocket. A common initial estimate for this residual propellant is five percent. With this ratio and

1520-406: Is achieved. In some cases with serial staging, the upper stage ignites before the separation—the interstage ring is designed with this in mind, and the thrust is used to help positively separate the two vehicles. Only multistage rockets have reached orbital speed . Single-stage-to-orbit designs are sought, but have not yet been demonstrated. Multi-stage rockets overcome a limitation imposed by

1596-453: Is generally assembled at its manufacturing site and shipped to the launch site; the term vehicle assembly refers to the mating of all rocket stage(s) and the spacecraft payload into a single assembly known as a space vehicle . Single-stage vehicles ( suborbital ), and multistage vehicles on the smaller end of the size range, can usually be assembled directly on the launch pad by lifting the stage(s) and spacecraft vertically in place by means of

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1672-577: Is impractical to directly compare the rocket's certain trait with the same trait of another because their individual attributes are often not independent of one another. For this reason, dimensionless ratios have been designed to enable a more meaningful comparison between rockets. The first is the initial to final mass ratio, which is the ratio between the rocket stage's full initial mass and the rocket stage's final mass once all of its fuel has been consumed. The equation for this ratio is: Where m E {\displaystyle m_{\mathrm {E} }}

1748-428: Is intermediate between a five-stage-to-orbit launcher and a three-stage-to-orbit launcher, most often used with solid-propellant launch systems. Other designs do not have all four stages inline on the main stack, instead having strap-on boosters for the "stage-0" with three core stages. In these designs, the boosters and first stage fire simultaneously instead of consecutively, providing extra initial thrust to lift

1824-496: Is only slightly less than on the Earth's surface. This is because the distance to LEO from the Earth's surface is much less than the Earth's radius. However, an object in orbit is in a permanent free fall around Earth, because in orbit the gravitational force and the centrifugal force balance each other out. As a result, spacecraft in orbit continue to stay in orbit, and people inside or outside such craft continuously experience weightlessness . Objects in LEO orbit Earth between

1900-514: Is the empty mass of the stage, m p {\displaystyle m_{\mathrm {p} }} is the mass of the propellant, and m P L {\displaystyle m_{\mathrm {PL} }} is the mass of the payload. The second dimensionless performance quantity is the structural ratio, which is the ratio between the empty mass of the stage, and the combined empty mass and propellant mass as shown in this equation: The last major dimensionless performance quantity

1976-402: Is the mass of the oxidizer and m f u e l {\displaystyle m_{\mathrm {fuel} }} is the mass of the fuel. This mixture ratio not only governs the size of each tank, but also the specific impulse of the rocket. Determining the ideal mixture ratio is a balance of compromises between various aspects of the rocket being designed, and can vary depending on

2052-455: Is the payload ratio, which is the ratio between the payload mass and the combined mass of the empty rocket stage and the propellant: After comparing the three equations for the dimensionless quantities, it is easy to see that they are not independent of each other, and in fact, the initial to final mass ratio can be rewritten in terms of structural ratio and payload ratio: These performance ratios can also be used as references for how efficient

2128-480: The Falcon 9 Full Thrust , are typically used to separate rocket stages. A two-stage-to-orbit ( TSTO ) or two-stage rocket launch vehicle is a spacecraft in which two distinct stages provide propulsion consecutively in order to achieve orbital velocity. It is intermediate between a three-stage-to-orbit launcher and a hypothetical single-stage-to-orbit (SSTO) launcher. The three-stage-to-orbit launch system

2204-734: The RTV-G-4 Bumper rockets tested at the White Sands Proving Ground and later at Cape Canaveral from 1948 to 1950. These consisted of a V-2 rocket and a WAC Corporal sounding rocket. The greatest altitude ever reached was 393 km, attained on February 24, 1949, at White Sands. In 1947, the Soviet rocket engineer and scientist Mikhail Tikhonravov developed a theory of parallel stages, which he called "packet rockets". In his scheme, three parallel stages were fired from liftoff , but all three engines were fueled from

2280-509: The Singijeon , or 'magical machine arrows' in the 16th century. The earliest experiments with multistage rockets in Europe were made in 1551 by Austrian Conrad Haas (1509–1576), the arsenal master of the town of Hermannstadt , Transylvania (now Sibiu/Hermannstadt, Romania). This concept was developed independently by at least five individuals: The first high-speed multistage rockets were

2356-490: The Soviet and U.S. space programs, were not passivated after mission completion. During the initial attempts to characterize the space debris problem, it became evident that a good proportion of all debris was due to the breaking up of rocket upper stages, particularly unpassivated upper-stage propulsion units. An illustration and description in the 14th century Chinese Huolongjing by Jiao Yu and Liu Bowen shows

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2432-399: The first stage is at the bottom and is usually the largest, the second stage and subsequent upper stages are above it, usually decreasing in size. In parallel staging schemes solid or liquid rocket boosters are used to assist with launch. These are sometimes referred to as "stage 0". In the typical case, the first-stage and booster engines fire to propel the entire rocket upwards. When

2508-523: The "stage-0" with two core stages. In these designs, the boosters and first stage fire simultaneously instead of consecutively, providing extra initial thrust to lift the full launcher weight and overcome gravity losses and atmospheric drag. The boosters are jettisoned a few minutes into flight to reduce weight. The four-stage-to-orbit launch system is a rocket system used to attain Earth orbit. The spacecraft uses four distinct stages to provide propulsion consecutively in order to achieve orbital velocity. It

2584-557: The LEO region near their lowest altitude (or perigee ) but are not in a LEO orbit because their highest altitude (or apogee ) exceeds 2,000 km (1,243 mi). Sub-orbital objects can also reach the LEO region but are not in a LEO orbit because they re-enter the atmosphere . The distinction between LEO orbits and the LEO region is especially important for analysis of possible collisions between objects which may not themselves be in LEO but could collide with satellites or debris in LEO orbits. The mean orbital velocity needed to maintain

2660-601: The Transfer Orbit Stage was an Orbus 21 solid rocket motor . Attitude control was provided by hydrazine thrusters. The inertial guidance system used a ring laser gyroscope produced by Honeywell . Its control system employed a digital optimal, time-scheduled control algorithm to provide stability whilst the Orbus 21 was burning, while a phase-plane controller was used to manage its reaction-control system. Only two Transfer Orbit Stages were launched. The first

2736-450: The anomaly. Upper stage By jettisoning stages when they run out of propellant, the mass of the remaining rocket is decreased. Each successive stage can also be optimized for its specific operating conditions, such as decreased atmospheric pressure at higher altitudes. This staging allows the thrust of the remaining stages to more easily accelerate the rocket to its final velocity and height. In serial or tandem staging schemes,

2812-462: The atmosphere and the exhaust gas does not need to expand against as much atmospheric pressure. When selecting the ideal rocket engine to use as an initial stage for a launch vehicle, a useful performance metric to examine is the thrust-to-weight ratio, and is calculated by the equation: The common thrust-to-weight ratio of a launch vehicle is within the range of 1.3 to 2.0. Another performance metric to keep in mind when designing each rocket stage in

2888-433: The atmosphere. The effects of adding such quantities of vaporized metals to Earth's stratosphere are potentially of concern but currently unknown. The LEO environment is becoming congested with space debris because of the frequency of object launches. This has caused growing concern in recent years, since collisions at orbital velocities can be dangerous or deadly. Collisions can produce additional space debris, creating

2964-403: The boosters run out of fuel, they are detached from the rest of the rocket (usually with some kind of small explosive charge or explosive bolts ) and fall away. The first stage then burns to completion and falls off. This leaves a smaller rocket, with the second stage on the bottom, which then fires. Known in rocketry circles as staging , this process is repeated until the desired final velocity

3040-401: The breakup of a single upper stage while in orbit. After the 1990s, spent upper stages are generally passivated after their use as a launch vehicle is complete in order to minimize risks while the stage remains derelict in orbit . Passivation means removing any sources of stored energy remaining on the vehicle, as by dumping fuel or discharging batteries. Many early upper stages, in both

3116-455: The denser part of the atmosphere and below the inner Van Allen radiation belt . They encounter atmospheric drag from gases in the thermosphere (approximately 80–600 km above the surface) or exosphere (approximately 600 km or 400 mi and higher), depending on orbit height. Satellites in orbits that reach altitudes below 300 km (190 mi) decay quickly due to atmospheric drag. Equatorial low Earth orbits ( ELEO ) are

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3192-419: The different stages of the rocket should be clearly defined. Continuing with the previous example, the end of the first stage which is sometimes referred to as 'stage 0', can be defined as when the side boosters separate from the main rocket. From there, the final mass of stage one can be considered the sum of the empty mass of stage one, the mass of stage two (the main rocket and the remaining unburned fuel) and

3268-404: The drawbacks of a less efficient specific impulse rating. But suppose the defining constraint for the launch system is volume, and a low density fuel is required such as hydrogen. This example would be solved by using an oxidizer-rich mixture ratio, reducing efficiency and specific impulse rating, but will meet a smaller tank volume requirement. The ultimate goal of optimal staging is to maximize

3344-400: The end of the rocket stage's motion, as the vehicle will still have a velocity that will allow it to coast upward for a brief amount of time until the acceleration of the planet's gravity gradually changes it to a downward direction. The velocity and altitude of the rocket after burnout can be easily modeled using the basic physics equations of motion. When comparing one rocket with another, it

3420-598: The entire vehicle stack to the launch pad in an upright position. In contrast, vehicles such as the Russian Soyuz rocket and the SpaceX Falcon 9 are assembled horizontally in a processing hangar, transported horizontally, and then brought upright at the pad. Spent upper stages of launch vehicles are a significant source of space debris remaining in orbit in a non-operational state for many years after use, and occasionally, large debris fields created from

3496-411: The equations for determining the burnout velocities, burnout times, burnout altitudes, and mass of each stage. This would make for a better approach to a conceptual design in a situation where a basic understanding of the system behavior is preferential to a detailed, accurate design. One important concept to understand when undergoing restricted rocket staging, is how the burnout velocity is affected by

3572-495: The farthest in LEO, before medium Earth orbit (MEO), have an altitude of 2,000 kilometers, about one-third of the radius of Earth and near the beginning of the inner Van Allen radiation belt . The term LEO region is used for the area of space below an altitude of 2,000 km (1,200 mi) (about one-third of Earth's radius). Objects in orbits that pass through this zone, even if they have an apogee further out or are sub-orbital , are carefully tracked since they present

3648-647: The first generation of Starlink satellites used polar orbits which provide coverage everywhere on Earth. Later Starlink constellations orbit at a lower inclination and provide more coverage for populated areas. Higher orbits include medium Earth orbit (MEO), sometimes called intermediate circular orbit (ICO), and further above, geostationary orbit (GEO). Orbits higher than low orbit can lead to early failure of electronic components due to intense radiation and charge accumulation. In 2017, " very low Earth orbits " ( VLEO ) began to be seen in regulatory filings. These orbits, below about 450 km (280 mi), require

3724-486: The first stage of the American Atlas I and Atlas II launch vehicles, arranged in a row, used parallel staging in a similar way: the outer pair of booster engines existed as a jettisonable pair which would, after they shut down, drop away with the lowermost outer skirt structure, leaving the central sustainer engine to complete the first stage's engine burn towards apogee or orbit. Separation of each portion of

3800-467: The fuel to be calculated if the density of the fuel is known, which is almost always the case when designing the rocket stage. The volume is yielded when dividing the mass of the propellant by its density. Asides from the fuel required, the mass of the rocket structure itself must also be determined, which requires taking into account the mass of the required thrusters, electronics, instruments, power equipment, etc. These are known quantities for typical off

3876-479: The full launcher weight and overcome gravity losses and atmospheric drag. The boosters are jettisoned a few minutes into flight to reduce weight. Low Earth orbit A low Earth orbit ( LEO ) is an orbit around Earth with a period of 128 minutes or less (making at least 11.25 orbits per day) and an eccentricity less than 0.25. Most of the artificial objects in outer space are in LEO, peaking in number at an altitude around 800 km (500 mi), while

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3952-405: The initial and final masses of the rocket stage respectively. In conjunction with the burnout time, the burnout height and velocity are obtained using the same values, and are found by these two equations: When dealing with the problem of calculating the total burnout velocity or time for the entire rocket system, the general procedure for doing so is as follows: The burnout time does not define

4028-443: The largest rocket ever to do so, as well as the first reusable vehicle to utilize hot staging. A rocket system that implements tandem staging means that each individual stage runs in order one after the other. The rocket breaks free from the previous stage, then begins burning through the next stage in straight succession. On the other hand, a rocket that implements parallel staging has two or more different stages that are active at

4104-429: The laws of physics on the velocity change achievable by a rocket stage. The limit depends on the fueled-to-dry mass ratio and on the effective exhaust velocity of the engine. This relation is given by the classical rocket equation : where: The delta v required to reach low Earth orbit (or the required velocity of a sufficiently heavy suborbital payload) requires a wet to dry mass ratio larger than has been achieved in

4180-403: The mass of the payload. High-altitude and space-bound upper stages are designed to operate with little or no atmospheric pressure. This allows the use of lower pressure combustion chambers and engine nozzles with optimal vacuum expansion ratios . Some upper stages, especially those using hypergolic propellants like Delta-K or Ariane 5 ES second stage, are pressure fed , which eliminates

4256-464: The mass of the propellant calculated, the mass of the empty rocket weight can be determined. Sizing rockets using a liquid bipropellant requires a slightly more involved approach because there are two separate tanks that are required: one for the fuel, and one for the oxidizer. The ratio of these two quantities is known as the mixture ratio, and is defined by the equation: Where m o x {\displaystyle m_{\mathrm {ox} }}

4332-515: The nearly spent stage keeps the propellants settled at the bottom of the tanks. Hot-staging is used on Soviet-era Russian rockets such as Soyuz and Proton-M . The N1 rocket was designed to use hot staging, however none of the test flights lasted long enough for this to occur. Starting with the Titan II, the Titan family of rockets used hot staging. SpaceX retrofitted their Starship rocket to use hot staging after its first flight , making it

4408-708: The need for complex turbopumps . Other upper stages, such as the Centaur or DCSS , use liquid hydrogen expander cycle engines, or gas generator cycle engines like the Ariane 5 ECA's HM7B or the S-IVB 's J-2 . These stages are usually tasked with completing orbital injection and accelerating payloads into higher energy orbits such as GTO or to escape velocity . Upper stages, such as Fregat , used primarily to bring payloads from low Earth orbit to GTO or beyond are sometimes referred to as space tugs . Each individual stage

4484-419: The number of stages that split up the rocket system. Increasing the number of stages for a rocket while keeping the specific impulse, payload ratios and structural ratios constant will always yield a higher burnout velocity than the same systems that use fewer stages. However, the law of diminishing returns is evident in that each increment in number of stages gives less of an improvement in burnout velocity than

4560-454: The oldest known multistage rocket; this was the " fire-dragon issuing from the water " (火龙出水, huǒ lóng chū shuǐ), which was used mostly by the Chinese navy. It was a two-stage rocket that had booster rockets that would eventually burn out, yet, before they did so, automatically ignited a number of smaller rocket arrows that were shot out of the front end of the missile, which was shaped like

4636-426: The orbit. Even for circular orbits , the altitude above ground can vary by as much as 30 km (19 mi) (especially for polar orbits ) due to the oblateness of Earth's spheroid figure and local topography . While definitions based on altitude are inherently ambiguous, most of them fall within the range specified by an orbit period of 128 minutes because, according to Kepler's third law , this corresponds to

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4712-525: The outer two stages, until they are empty and could be ejected. This is more efficient than sequential staging, because the second-stage engine is never just dead weight. In 1951, Soviet engineer and scientist Dmitry Okhotsimsky carried out a pioneering engineering study of general sequential and parallel staging, with and without the pumping of fuel between stages. The design of the R-7 Semyorka emerged from that study. The trio of rocket engines used in

4788-632: The overall payload ratio of the entire system. It is important to note that when computing payload ratio for individual stages, the payload includes the mass of all the stages after the current one. The overall payload ratio is: Where n is the number of stages the rocket system comprises. Similar stages yielding the same payload ratio simplify this equation, however that is seldom the ideal solution for maximizing payload ratio, and ΔV requirements may have to be partitioned unevenly as suggested in guideline tips 1 and 2 from above. Two common methods of determining this perfect ΔV partition between stages are either

4864-573: The payload ratio (see ratios under performance), meaning the largest amount of payload is carried up to the required burnout velocity using the least amount of non-payload mass, which comprises everything else. This goal assumes that the cost of a rocket launch is proportional to the total liftoff mass of the rocket, which is a rule of thumb in rocket engineering. Here are a few quick rules and guidelines to follow in order to reach optimal staging: The payload ratio can be calculated for each individual stage, and when multiplied together in sequence, will yield

4940-402: The previous increment. The burnout velocity gradually converges towards an asymptotic value as the number of stages increases towards a very high number. In addition to diminishing returns in burnout velocity improvement, the main reason why real world rockets seldom use more than three stages is because of increase of weight and complexity in the system for each added stage, ultimately yielding

5016-427: The rocket system. Restricted rocket staging is based on the simplified assumption that each of the stages of the rocket system have the same specific impulse, structural ratio, and payload ratio, the only difference being the total mass of each increasing stage is less than that of the previous stage. Although this assumption may not be the ideal approach to yielding an efficient or optimal system, it greatly simplifies

5092-518: The same time. For example, the Space Shuttle has two Solid Rocket Boosters that burn simultaneously. Upon launch, the boosters ignite, and at the end of the stage, the two boosters are discarded while the external fuel tank is kept for another stage. Most quantitative approaches to the design of the rocket system's performance are focused on tandem staging, but the approach can be easily modified to include parallel staging. To begin with,

5168-416: The savings are so great that every rocket ever used to deliver a payload into orbit has had staging of some sort. One of the most common measures of rocket efficiency is its specific impulse, which is defined as the thrust per flow rate (per second) of propellant consumption: When rearranging the equation such that thrust is calculated as a result of the other factors, we have: These equations show that

5244-518: The shelf hardware that should be considered in the mid to late stages of the design, but for preliminary and conceptual design, a simpler approach can be taken. Assuming one engine for a rocket stage provides all of the total impulse for that particular segment, a mass fraction can be used to determine the mass of the system. The mass of the stage transfer hardware such as initiators and safe-and-arm devices are very small by comparison and can be considered negligible. For modern day solid rocket motors, it

5320-414: The spent lower stages. A further advantage is that each stage can use a different type of rocket engine, each tuned for its particular operating conditions. Thus the lower-stage engines are designed for use at atmospheric pressure, while the upper stages can use engines suited to near vacuum conditions. Lower stages tend to require more structure than upper as they need to bear their own weight plus that of

5396-486: The stages above them. Optimizing the structure of each stage decreases the weight of the total vehicle and provides further advantage. The advantage of staging comes at the cost of the lower stages lifting engines which are not yet being used, as well as making the entire rocket more complex and harder to build than a single stage. In addition, each staging event is a possible point of launch failure, due to separation failure, ignition failure, or stage collision. Nevertheless,

5472-406: The type of fuel and oxidizer combination being used. For example, a mixture ratio of a bipropellant could be adjusted such that it may not have the optimal specific impulse, but will result in fuel tanks of equal size. This would yield simpler and cheaper manufacturing, packing, configuring, and integrating of the fuel systems with the rest of the rocket, and can become a benefit that could outweigh

5548-412: The use of novel technologies for orbit raising because they operate in orbits that would ordinarily decay too soon to be economically useful. A low Earth orbit requires the lowest amount of energy for satellite placement. It provides high bandwidth and low communication latency . Satellites and space stations in LEO are more accessible for crew and servicing. Since it requires less energy to place

5624-656: Was designed to separate from the Space Shuttle via pyrotechnic devices enclosed in a frangible metal joint known as a Super*Zip. On STS-51, the primary and backup pyrotechnic devices inside the Super*Zip were erroneously fired simultaneously, resulting in minor damage to both the TOS, and the Shuttle's payload bay lining. However, the ACTS satellite deployed successfully and functioned normally on orbit. The Shuttle landed safely despite

5700-637: Was launched on 25 September 1992, aboard a Commercial Titan III with the Mars Observer spacecraft bound for Mars . Despite a successful launch, the Mars Observer spacecraft later malfunctioned. The second use of a TOS was the deployment of the Advanced Communications Technology Satellite from Space Shuttle Discovery on mission STS-51 . This was also successful. The Transfer Orbit Stage

5776-515: Was proposed by medieval Korean engineer, scientist and inventor Ch'oe Mu-sŏn and developed by the Firearms Bureau (火㷁道監) during the 14th century. The rocket had the length of 15 cm and 13 cm; the diameter was 2.2 cm. It was attached to an arrow 110 cm long; experimental records show that the first results were around 200m in range. There are records that show Korea kept developing this technology until it came to produce

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