Orbital maneuver

From Wikipedia, the free encyclopedia

https://en.wikipedia.org/wiki/Orbital_maneuver 

In spaceflight, an orbital maneuver (otherwise known as a burn) is the use of propulsion systems to change the orbit of a spacecraft. For spacecraft far from Earth (for example those in orbits around the Sun) an orbital maneuver is called a deep-space maneuver (DSM).

The rest of the flight, especially in a transfer orbit, is called coasting.

General

Rocket equation

Main article: Tsiolkovsky rocket equation

Rocket mass ratios versus final velocity calculated from the rocket equation

The Tsiolkovsky rocket equation, or ideal rocket equation is an equation that is useful for considering vehicles that follow the basic principle of a rocket: where a device that can apply acceleration to itself (a thrust) by expelling part of its mass with high speed and moving due to the conservation of momentum. Specifically, it is a mathematical equation that relates the delta-v (the maximum change of speed of the rocket if no other external forces act) with the effective exhaust velocity and the initial and final mass of a rocket (or other reaction engine.)

For any such maneuver (or journey involving a number of such maneuvers):

${\displaystyle \mathrm{\Delta }v=v_{\text{e}}\ln \frac{m_0}{m_1}}$

where:

${\displaystyle m_0}$ is the initial total mass, including propellant,

${\displaystyle m_1}$ is the final total mass,

${\displaystyle v_{\text{e}}}$ is the effective exhaust velocity (${\displaystyle v_{\text{e}}=I_{\text{sp}}\cdot g_0}$ where ${\displaystyle I_{\text{sp}}}$ is the specific impulse expressed as a time period and ${\displaystyle g_0}$ is standard gravity),

${\displaystyle \mathrm{\Delta }v}$ is delta-v - the maximum change of velocity of the vehicle (with no external forces acting).

Delta-v

Main articles: Delta-v and delta-v budget

The applied change in velocity of each maneuver is referred to as delta-v (${\displaystyle \mathrm{\Delta }\mathbf{v}}$).

The delta-v for all the expected maneuvers are estimated for a mission. They are summarized in a delta-v budget. With a good approximation of the delta-v budget designers can estimate the fuel to payload requirements of the spacecraft using the rocket equation.

Impulsive maneuvers

Figure 1: Approximation of a finite thrust maneuver with an impulsive change in velocity

An "impulsive maneuver" is the mathematical model of a maneuver as an instantaneous change in the spacecraft's velocity (magnitude and/or direction) as illustrated in figure 1. It is the limit case of a burn to generate a particular amount of delta-v, as the burn time tends to zero.

In the physical world no truly instantaneous change in velocity is possible as this would require an "infinite force" applied during an "infinitely short time" but as a mathematical model it in most cases describes the effect of a maneuver on the orbit very well.

The off-set of the velocity vector after the end of real burn from the velocity vector at the same time resulting from the theoretical impulsive maneuver is only caused by the difference in gravitational force along the two paths (red and black in figure 1) which in general is small.

In the planning phase of space missions designers will first approximate their intended orbital changes using impulsive maneuvers that greatly reduces the complexity of finding the correct orbital transitions.

Low thrust for a long time

Applying a low thrust over a longer period of time is referred to as a non-impulsive maneuver. 'Non-impulsive' refers to the momentum changing slowly over a long time, as in electrically powered spacecraft propulsion, rather than by a short impulse.

Another term is finite burn, where the word "finite" is used to mean "non-zero", or practically, again: over a longer period.

For a few space missions, such as those including a space rendezvous, high fidelity models of the trajectories are required to meet the mission goals. Calculating a "finite" burn requires a detailed model of the spacecraft and its thrusters. The most important of details include: mass, center of mass, moment of inertia, thruster positions, thrust vectors, thrust curves, specific impulse, thrust centroid offsets, and fuel consumption.

Assists

Oberth effect

Main article: Oberth effect

In astronautics, the Oberth effect is where the use of a rocket engine when travelling at high speed generates much more useful energy than one at low speed. Oberth effect occurs because the propellant has more usable energy (due to its kinetic energy on top of its chemical potential energy) and it turns out that the vehicle is able to employ this kinetic energy to generate more mechanical power. It is named after Hermann Oberth, the Austro-Hungarian-born, German physicist and a founder of modern rocketry, who apparently first described the effect.

The Oberth effect is used in a powered flyby or Oberth maneuver where the application of an impulse, typically from the use of a rocket engine, close to a gravitational body (where the gravity potential is low, and the speed is high) can give much more change in kinetic energy and final speed (i.e. higher specific energy) than the same impulse applied further from the body for the same initial orbit.

Since the Oberth maneuver happens in a very limited time (while still at low altitude), to generate a high impulse the engine necessarily needs to achieve high thrust (impulse is by definition the time multiplied by thrust). Thus the Oberth effect is far less useful for low-thrust engines, such as ion thrusters.

Historically, a lack of understanding of this effect led investigators to conclude that interplanetary travel would require completely impractical amounts of propellant, as without it, enormous amounts of energy are needed.

Gravitational assist

Main article: Gravity assist

The trajectories that enabled NASA's twin Voyager spacecraft to tour the four gas giant planets and achieve velocity to escape our solar system

In orbital mechanics and aerospace engineering, a gravitational slingshot, gravity assist maneuver, or swing-by is the use of the relative movement and gravity of a planet or other celestial body to alter the path and speed of a spacecraft, typically in order to save propellant, time, and expense. Gravity assistance can be used to accelerate, decelerate and/or re-direct the path of a spacecraft.

The "assist" is provided by the motion (orbital angular momentum) of the gravitating body as it pulls on the spacecraft. The technique was first proposed as a mid-course manoeuvre in 1961, and used by interplanetary probes from Mariner 10 onwards, including the two Voyager probes' notable fly-bys of Jupiter and Saturn.

Transfer orbits

Orbit insertion is a general term for a maneuver that is more than a small correction. It may be used for a maneuver to change a transfer orbit or an ascent orbit into a stable one, but also to change a stable orbit into a descent: descent orbit insertion. Also the term orbit injection is used, especially for changing a stable orbit into a transfer orbit, e.g. trans-lunar injection (TLI), trans-Mars injection (TMI) and trans-Earth injection (TEI).

Hohmann transfer

Hohmann Transfer Orbit

Main article: Hohmann transfer orbit

In orbital mechanics, the Hohmann transfer orbit is an elliptical orbit used to transfer between two circular orbits of different altitudes, in the same plane.

The orbital maneuver to perform the Hohmann transfer uses two engine impulses which move a spacecraft onto and off the transfer orbit. This maneuver was named after Walter Hohmann, the German scientist who published a description of it in his 1925 book Die Erreichbarkeit der Himmelskörper (The Accessibility of Celestial Bodies). Hohmann was influenced in part by the German science fiction author Kurd Laßwitz and his 1897 book Two Planets.

Bi-elliptic transfer

Bi-elliptic transfer from blue to red circular orbit

Main article: Bi-elliptic transfer

In astronautics and aerospace engineering, the bi-elliptic transfer is an orbital maneuver that moves a spacecraft from one orbit to another and may, in certain situations, require less delta-v than a Hohmann transfer maneuver.

The bi-elliptic transfer consists of two half elliptic orbits. From the initial orbit, a delta-v is applied boosting the spacecraft into the first transfer orbit with an apoapsis at some point ${\displaystyle r_b}$ away from the central body. At this point, a second delta-v is applied sending the spacecraft into the second elliptical orbit with periapsis at the radius of the final desired orbit, where a third delta-v is performed, injecting the spacecraft into the desired orbit.

While they require one more engine burn than a Hohmann transfer and generally requires a greater travel time, some bi-elliptic transfers require a lower amount of total delta-v than a Hohmann transfer when the ratio of final to initial semi-major axis is 11.94 or greater, depending on the intermediate semi-major axis chosen.

The idea of the bi-elliptical transfer trajectory was first published by Ary Sternfeld in 1934.

Low energy transfer

Main article: low energy transfer

A low energy transfer, or low energy trajectory, is a route in space which allows spacecraft to change orbits using very little fuel. These routes work in the Earth-Moon system and also in other systems, such as traveling between the satellites of Jupiter. The drawback of such trajectories is that they take much longer to complete than higher energy (more fuel) transfers such as Hohmann transfer orbits.

Low energy transfer are also known as weak stability boundary trajectories, or ballistic capture trajectories.

Low energy transfers follow special pathways in space, sometimes referred to as the Interplanetary Transport Network. Following these pathways allows for long distances to be traversed for little expenditure of delta-v.

Orbital inclination change

Main article: orbital inclination change

Orbital inclination change is an orbital maneuver aimed at changing the inclination of an orbiting body's orbit. This maneuver is also known as an orbital plane change as the plane of the orbit is tipped. This maneuver requires a change in the orbital velocity vector (delta v) at the orbital nodes (i.e. the point where the initial and desired orbits intersect, the line of orbital nodes is defined by the intersection of the two orbital planes).

In general, inclination changes can require a great deal of delta-v to perform, and most mission planners try to avoid them whenever possible to conserve fuel. This is typically achieved by launching a spacecraft directly into the desired inclination, or as close to it as possible so as to minimize any inclination change required over the duration of the spacecraft life.

Maximum efficiency of inclination change is achieved at apoapsis, (or apogee), where orbital velocity ${\displaystyle v}$ is the lowest. In some cases, it may require less total delta v to raise the spacecraft into a higher orbit, change the orbit plane at the higher apogee, and then lower the spacecraft to its original altitude.

Constant-thrust trajectory

Constant-thrust and constant-acceleration trajectories involve the spacecraft firing its engine in a prolonged constant burn. In the limiting case where the vehicle acceleration is high compared to the local gravitational acceleration, the spacecraft points straight toward the target (accounting for target motion), and remains accelerating constantly under high thrust until it reaches its target. In this high-thrust case, the trajectory approaches a straight line. If it is required that the spacecraft rendezvous with the target, rather than performing a flyby, then the spacecraft must flip its orientation halfway through the journey, and decelerate the rest of the way.

In the constant-thrust trajectory, the vehicle's acceleration increases during thrusting period, since the fuel use means the vehicle mass decreases. If, instead of constant thrust, the vehicle has constant acceleration, the engine thrust must decrease during the trajectory.

This trajectory requires that the spacecraft maintain a high acceleration for long durations. For interplanetary transfers, days, weeks or months of constant thrusting may be required. As a result, there are no currently available spacecraft propulsion systems capable of using this trajectory. It has been suggested that some forms of nuclear (fission or fusion based) or antimatter powered rockets would be capable of this trajectory.

More practically, this type of maneuver is used in low thrust maneuvers, for example with ion engines, Hall-effect thrusters, and others. These types of engines have very high specific impulse (fuel efficiency) but currently are only available with fairly low absolute thrust.

Rendezvous and docking

Orbit phasing

Main article: Orbit phasing

In astrodynamics orbit phasing is the adjustment of the time-position of spacecraft along its orbit, usually described as adjusting the orbiting spacecraft's true anomaly.

Space rendezvous and docking

Main article: Space rendezvous

Gemini 7 photographed from Gemini 6 in December 1965

A space rendezvous is an orbital maneuver during which two spacecraft, one of which is often a space station, arrive at the same orbit and approach to a very close distance (e.g. within visual contact). Rendezvous requires a precise match of the orbital velocities of the two spacecraft, allowing them to remain at a constant distance through orbital station-keeping. Rendezvous may or may not be followed by docking or berthing, procedures which bring the spacecraft into physical contact and create a link between them.

Delta-v

From Wikipedia, the free encyclopedia

https://en.wikipedia.org/wiki/Delta-v

Delta-v (more known as "change in velocity"), symbolized as $\mathrm{\Delta }v$ and pronounced delta-vee, as used in spacecraft flight dynamics, is a measure of the impulse per unit of spacecraft mass that is needed to perform a maneuver such as launching from or landing on a planet or moon, or an in-space orbital maneuver. It is a scalar that has the units of speed. As used in this context, it is not the same as the physical change in velocity of said spacecraft.

A simple example might be the case of a conventional rocket-propelled spacecraft, which achieves thrust by burning fuel. Such a spacecraft's delta-v, then, would be the change in velocity that spacecraft can achieve by burning its entire fuel load.

Delta-v is produced by reaction engines, such as rocket engines, and is proportional to the thrust per unit mass and the burn time. It is used to determine the mass of propellant required for the given maneuver through the Tsiolkovsky rocket equation.

For multiple maneuvers, delta-v sums linearly.

For interplanetary missions, delta-v is often plotted on a porkchop plot, which displays the required mission delta-v as a function of launch date.

Definition

$${\displaystyle \mathrm{\Delta }v={\int }_{t_0}^{t_1}\frac{|T(t)|}{m(t)}dt}$$

where

• T(t) is the instantaneous thrust at time t.
• m(t) is the instantaneous mass at time t.

Specific cases

In the absence of external forces:

$${\displaystyle \mathrm{\Delta }v={\int }_{t_0}^{t_1}\left|\dot{v}\right|dt}$$

where ${\displaystyle \dot{v}}$ is the coordinate acceleration.

When thrust is applied in a constant direction (v/|v| is constant) this simplifies to:

$${\displaystyle \mathrm{\Delta }v=|v_1-v_0|}$$

which is simply the magnitude of the change in velocity. However, this relation does not hold in the general case: if, for instance, a constant, unidirectional acceleration is reversed after (t₁ − t₀)/2 then the velocity difference is 0, but delta-v is the same as for the non-reversed thrust.

For rockets, "absence of external forces" is taken to mean the absence of gravity and atmospheric drag, as well as the absence of aerostatic back pressure on the nozzle, and hence the vacuum I_sp is used for calculating the vehicle's delta-v capacity via the rocket equation. In addition, the costs for atmospheric losses and gravity drag are added into the delta-v budget when dealing with launches from a planetary surface.

Orbital maneuvers

Main article: rocket equation

Orbit maneuvers are made by firing a thruster to produce a reaction force acting on the spacecraft. The size of this force will be

${\displaystyle T=v_{\text{exh}}\rho }$

(1)

where

• v_exh is the velocity of the exhaust gas in rocket frame
• ρ is the propellant flow rate to the combustion chamber

The acceleration ${\displaystyle \dot{v}}$ of the spacecraft caused by this force will be

${\displaystyle \dot{v}=\frac{T}{m}=v_{\text{exh}}\frac{\rho }{m}}$

(2)

where m is the mass of the spacecraft

During the burn the mass of the spacecraft will decrease due to use of fuel, the time derivative of the mass being

${\displaystyle \dot{m}=-\rho }$

(3)

If now the direction of the force, i.e. the direction of the nozzle, is fixed during the burn one gets the velocity increase from the thruster force of a burn starting at time ${\displaystyle t_0}$ and ending at t₁ as

${\displaystyle \mathrm{\Delta }v=-{\int }_{t_0}^{t_1}{v}_{\text{exh}}\frac{\dot{m}}{m}dt}$

(4)

Changing the integration variable from time t to the spacecraft mass m one gets

${\displaystyle \mathrm{\Delta }v=-{\int }_{m_0}^{m_1}{v}_{\text{exh}}\frac{dm}{m}}$

(5)

Assuming ${\displaystyle v_{\text{exh}}}$ to be a constant not depending on the amount of fuel left this relation is integrated to

${\displaystyle \mathrm{\Delta }v=v_{\text{exh}}\ln \left(\frac{m_0}{m_1}\right)}$

(6)

which is the Tsiolkovsky rocket equation.

If for example 20% of the launch mass is fuel giving a constant ${\displaystyle v_{\text{exh}}}$ of 2100 m/s (a typical value for a hydrazine thruster) the capacity of the reaction control system is

$${\displaystyle \mathrm{\Delta }v=2100\ln \left(\frac{1}{0.8}\right)\text{m/s}=460\text{m/s}.}$$

If ${\displaystyle v_{\text{exh}}}$ is a non-constant function of the amount of fuel left

$${\displaystyle v_{\text{exh}}=v_{\text{exh}}(m)}$$

the capacity of the reaction control system is computed by the integral (5).

The acceleration (2) caused by the thruster force is just an additional acceleration to be added to the other accelerations (force per unit mass) affecting the spacecraft and the orbit can easily be propagated with a numerical algorithm including also this thruster force. But for many purposes, typically for studies or for maneuver optimization, they are approximated by impulsive maneuvers as illustrated in figure 1 with a ${\displaystyle \mathrm{\Delta }v}$ as given by (4). Like this one can for example use a "patched conics" approach modeling the maneuver as a shift from one Kepler orbit to another by an instantaneous change of the velocity vector.

Figure 1: Approximation of a finite thrust maneuver with an impulsive change in velocity having the delta-v given by (4).

This approximation with impulsive maneuvers is in most cases very accurate, at least when chemical propulsion is used. For low thrust systems, typically electrical propulsion systems, this approximation is less accurate. But even for geostationary spacecraft using electrical propulsion for out-of-plane control with thruster burn periods extending over several hours around the nodes this approximation is fair.

Production

Delta-v is typically provided by the thrust of a rocket engine, but can be created by other engines. The time-rate of change of delta-v is the magnitude of the acceleration caused by the engines, i.e., the thrust per total vehicle mass. The actual acceleration vector would be found by adding thrust per mass on to the gravity vector and the vectors representing any other forces acting on the object.

The total delta-v needed is a good starting point for early design decisions since consideration of the added complexities are deferred to later times in the design process.

The rocket equation shows that the required amount of propellant dramatically increases with increasing delta-v. Therefore, in modern spacecraft propulsion systems considerable study is put into reducing the total delta-v needed for a given spaceflight, as well as designing spacecraft that are capable of producing larger delta-v.

Increasing the delta-v provided by a propulsion system can be achieved by:

• staging
• increasing specific impulse
• improving propellant mass fraction

Multiple maneuvers

Because the mass ratios apply to any given burn, when multiple maneuvers are performed in sequence, the mass ratios multiply.

Thus it can be shown that, provided the exhaust velocity is fixed, this means that delta-v can be summed:

When m₁, m₂ are the mass ratios of the maneuvers, and v₁, v₂ are the delta-v of the first and second maneuvers

$${\displaystyle \begin{array}{rl}{m}_1{m}_2& =e^{V_1/V_e}{e}^{V_2/V_e}\\ & =e^{\frac{V_1+V_2}{V_e}}\\ & =e^{V/V_e}=M\end{array}}$$

where V = v₁ + v₂ and M = m₁ m₂. This is just the rocket equation applied to the sum of the two maneuvers.

This is convenient since it means that delta-v can be calculated and simply added and the mass ratio calculated only for the overall vehicle for the entire mission. Thus delta-v is commonly quoted rather than mass ratios which would require multiplication.

Delta-v budgets

Delta-v map of selected bodies in the solar system, assuming burns are at periapsis, and gravity assist and inclination changes are ignored

Main article: Delta-v budget

When designing a trajectory, delta-v budget is used as a good indicator of how much propellant will be required. Propellant usage is an exponential function of delta-v in accordance with the rocket equation, it will also depend on the exhaust velocity.

It is not possible to determine delta-v requirements from conservation of energy by considering only the total energy of the vehicle in the initial and final orbits since energy is carried away in the exhaust (see also below). For example, most spacecraft are launched in an orbit with inclination fairly near to the latitude at the launch site, to take advantage of the Earth's rotational surface speed. If it is necessary, for mission-based reasons, to put the spacecraft in an orbit of different inclination, a substantial delta-v is required, though the specific kinetic and potential energies in the final orbit and the initial orbit are equal.

When rocket thrust is applied in short bursts the other sources of acceleration may be negligible, and the magnitude of the velocity change of one burst may be simply approximated by the delta-v. The total delta-v to be applied can then simply be found by addition of each of the delta-v's needed at the discrete burns, even though between bursts the magnitude and direction of the velocity changes due to gravity, e.g. in an elliptic orbit.

For examples of calculating delta-v, see Hohmann transfer orbit, gravitational slingshot, and Interplanetary Transport Network. It is also notable that large thrust can reduce gravity drag.

Delta-v is also required to keep satellites in orbit and is expended in propulsive orbital stationkeeping maneuvers. Since the propellant load on most satellites cannot be replenished, the amount of propellant initially loaded on a satellite may well determine its useful lifetime.

Oberth effect

Main article: Oberth effect

From power considerations, it turns out that when applying delta-v in the direction of the velocity the specific orbital energy gained per unit delta-v is equal to the instantaneous speed. This is called the Oberth effect.

For example, a satellite in an elliptical orbit is boosted more efficiently at high speed (that is, small altitude) than at low speed (that is, high altitude).

Another example is that when a vehicle is making a pass of a planet, burning the propellant at closest approach rather than further out gives significantly higher final speed, and this is even more so when the planet is a large one with a deep gravity field, such as Jupiter.

See also powered slingshots.

Porkchop plot

Main article: Porkchop plot

Due to the relative positions of planets changing over time, different delta-vs are required at different launch dates. A diagram that shows the required delta-v plotted against time is sometimes called a porkchop plot. Such a diagram is useful since it enables calculation of a launch window, since launch should only occur when the mission is within the capabilities of the vehicle to be employed.

Around the Solar System

Delta-v needed for various orbital manoeuvers using conventional rockets; red arrows show where optional aerobraking can be performed in that particular direction, black numbers give delta-v in km/s that apply in either direction. Lower-delta-v transfers than shown can often be achieved, but involve rare transfer windows or take significantly longer, see: fuzzy orbital transfers.

C3

Escape orbit

GEO

Geosynchronous orbit

GTO

Geostationary transfer orbit

L4/5

Earth–Moon L₄L₅ Lagrangian point

LEO

Low Earth orbit

LEO reentry

For example the Soyuz spacecraft makes a de-orbit from the ISS in two steps. First, it needs a delta-v of 2.18 m/s for a safe separation from the space station. Then it needs another 128 m/s for reentry.

Saturn (rocket family)

From Wikipedia, the free encyclopedia

https://en.wikipedia.org/wiki/Saturn_(rocket_family) 

Three variants of the Saturn family which were developed: Saturn I, Saturn IB, and Saturn V

The Saturn family of American rockets was developed by a team of mostly Nazi rocket engineers and scientists led by Wernher von Braun to launch heavy payloads to Earth orbit and beyond. The Saturn family used liquid hydrogen as fuel in the upper stages. Originally proposed as a military satellite launcher, they were adopted as the launch vehicles for the Apollo Moon program. Three versions were built and flown: the medium-lift Saturn I, the heavy-lift Saturn IB, and the super heavy-lift Saturn V.

The Saturn name was proposed by von Braun in October 1958 as a logical successor to the Jupiter series as well as the Roman god's powerful position.

In 1963, President John F. Kennedy identified the Saturn I SA-5 launch as being the point where US lift capability would surpass the Soviets, after having been behind since Sputnik. He last mentioned this in a speech given at Brooks AFB in San Antonio on the day before he was assassinated.

To date, the Saturn V is the only launch vehicle to transport human beings beyond low Earth orbit. A total of 24 humans were flown to the Moon in the four years spanning December 1968 through December 1972. No Saturn rocket failed catastrophically in flight.

Summary of variants

All the Saturn family rockets are listed here by date of introduction.

Name	Serial number	Function	Maiden flight	Final flight	Launches			Remarks
					Total	Success	Failure (+ partial)
Saturn I Block I	SA–1–4	Development	October 27, 1961	March 28, 1963	4	4	0	Live first stage only
Saturn I Block II	SA–5–10	Development	January 29, 1964	July 30, 1965	6	6	0	Carried Apollo boilerplate CSM and Pegasus micrometeroid satellites.
Saturn IB	SA–200	Apollo spacecraft Earth orbital carrier	February 26, 1966	July 15, 1975	9	9	0	Used for Apollo 7, Skylab crews and Apollo-Soyuz Test Project
Saturn V	SA–500	Apollo spacecraft lunar carrier	November 9, 1967	May 14, 1973	13	12	1	Launched nine crewed lunar missions and the Skylab space station

History

Early development

A Saturn I (SA-1) liftoff from LC-34

In the early 1950s, the US Navy and US Army actively developed long-range missiles with the help of German rocket engineers who were involved in developing the successful V-2 during the Second World War. These missiles included the Navy's Viking, and the Army's Corporal, Jupiter and Redstone. Meanwhile, the US Air Force developed its Atlas and Titan missiles, relying more on American engineers.

Infighting among the various branches was constant, with the United States Department of Defense (DoD) deciding which projects to fund for development. On November 26, 1956, Defense Secretary Charles E. Wilson issued a memorandum stripping the Army of offensive missiles with a range of 200 miles (320 km) or greater, and turning their Jupiter missiles over to the Air Force. From that point on, the Air Force would be the primary missile developer, especially for dual-use missiles that could also be used as space launch vehicles.

In late 1956, the Department of Defense released a requirement for a heavy-lift vehicle to orbit a new class of communications and "other" satellites (the spy satellite program was top secret). The requirements, drawn up by the then-unofficial Advanced Research Projects Agency (ARPA), called for a vehicle capable of putting 9,000 to 18,000 kilograms into orbit, or accelerating 2,700 to 5,400 kg to escape velocity.

Since the Wilson memorandum covered only weapons, not space vehicles, the Army Ballistic Missile Agency (ABMA) saw this as a way to continue the development of their own large-rocket projects. In April 1957, von Braun directed Heinz-Hermann Koelle, chief of the Future Projects design branch, to study dedicated launch vehicle designs that could be built as quickly as possible. Koelle evaluated a variety of designs for missile-derived launchers that could place a maximum of about 1,400 kg in orbit, but might be expanded to as much as 4,500 kg with new high-energy upper stages. In any event, these upper stages would not be available until 1961 or 1962 at the earliest, and the launchers would still not meet the DoD requirements for heavy loads.

In order to fill the projected need for loads of 10,000 kg or greater, the ABMA team calculated that a booster (first stage) with a thrust of about 1,500,000 lbf (6,700 kN) thrust would be needed, far greater than any existing or planned missile. For this role they proposed using a number of existing missiles clustered together to produce a single larger booster; using existing designs they looked at combining tankage from one Jupiter as a central core, with eight Redstone diameter tanks attached to it. This relatively cheap configuration allowed existing fabrication and design facilities to be used to produce this "quick and dirty" design.

Two approaches to building the Super-Jupiter were considered; the first used multiple engines to reach the 1,500,000 lbf (6,700 kN) mark, the second used a single much larger engine. Both approaches had their own advantages and disadvantages. Building a smaller engine for clustered use would be a relatively low-risk path from existing systems, but required duplication of systems and made the possibility of a stage failure much higher (adding engines generally reduces reliability, as per Lusser's law). A single larger engine would be more reliable, and would offer higher performance because it eliminated duplication of "dead weight" like propellant plumbing and hydraulics for steering the engines. On the downside, an engine of this size had never been built before and development would be expensive and risky. The Air Force had recently expressed an interest in such an engine, which would develop into the famed F-1, but at the time they were aiming for 1,000,000 lbf (4,400 kN) and the engines would not be ready until the mid-1960s. The engine-cluster appeared to be the only way to meet the requirements on time and budget.

Super-Jupiter was the first-stage booster only; to place payloads in orbit, additional upper stages would be needed. ABMA proposed using either the Titan or Atlas as a second stage, optionally with the new Centaur upper-stage. The Centaur had been proposed by General Dynamics (Astronautics Corp.) as an upper stage for the Atlas (also their design) in order to quickly produce a launcher capable of placing loads up to 8,500 lb (3,900 kg) into low Earth orbit. The Centaur was based on the same "balloon tank" concept as the Atlas, and built on the same jigs at the same 120-inch (3,000 mm) diameter. As the Titan was deliberately built at the same size as well, this meant the Centaur could be used with either missile. Given that the Atlas was the higher priority of the two ICBM projects and its production was fully accounted for, ABMA focused on "backup" design, Titan, although they proposed extending it in length in order to carry additional fuel.

In December 1957, ABMA delivered Proposal: A National Integrated Missile and Space Vehicle Development Program to the DoD, detailing their clustered approach. They proposed a booster consisting of a Jupiter missile airframe surrounded by eight Redstones acting as tankage, a thrust plate at the bottom, and four Rocketdyne E-1 engines, each having 380,000 lbf (1,700 kN) of thrust. The ABMA team also left the design open to future expansion with a single 1,500,000 lbf (6,700 kN) engine, which would require relatively minor changes to the design. The upper stage was the lengthened Titan, with the Centaur on top. The result was a very tall and skinny rocket, quite different from the Saturn that eventually emerged.

Specific uses were forecast for each of the military services, including navigation satellites for the Navy; reconnaissance, communications, and meteorological satellites for the Army and Air Force; support for Air Force crewed missions; and surface-to-surface logistics supply for the Army at distances up to 6400 km. Development and testing of the lower stage stack were projected to be completed by 1963, about the same time that the Centaur should become available for testing in combination. The total development cost of $850 million during the years 1958-1963 covered 30 research and development flights.

Sputnik stuns the world

While the Super-Jupiter program was being drawn up, preparations were underway for the first satellite launch as the US contribution to the International Geophysical Year in 1957. For complex political reasons, the program had been given to the US Navy under Project Vanguard. The Vanguard launcher consisted of a Viking lower stage combined with new uppers adapted from sounding rockets. ABMA provided valuable support on Viking and Vanguard, both with their first-hand knowledge of the V-2, as well as developing its guidance system. The first three Vanguard suborbital test flights had gone off without a hitch, starting in December 1956, and a launch was planned for late 1957.

On October 4, 1957, the Soviet Union surprised the world with the launch of Sputnik I. Although there had been some indications that the Soviets were working towards this goal, few in the U.S. military and scientific establishment considered these efforts seriously.

When asked in November 1954 about the possibility of the Soviets launching a satellite, Defense Secretary Wilson replied: "I wouldn't care if they did." The public did not see it the same way, however, and the event was a major public relations disaster for the US. Vanguard was planned to launch shortly after Sputnik, but a series of delays pushed this into December, when the rocket exploded in spectacular fashion. The press was harsh, referring to the project as "Kaputnik" or "Project Rearguard". As Time magazine noted at the time:

But in the midst of the cold war, Vanguard's cool scientific goal proved to be disastrously modest: the Russians got there first. The post-Sputnik White House explanation that the U.S. was not in a satellite "race" with Russia was not just an after-the-fact alibi. Said Dr. Hagen ten months ago: "We are not attempting in any way to race with the Russians". But in the eyes of the world, the U.S. was in a satellite race whether it wanted to be or not, and because of the Administration's costly failure of imagination, Project Vanguard shuffled along when it should have been running. It was still shuffling when Sputnik's beeps told the world that Russia's satellite program, not the U.S.'s, was the vanguard.

Von Braun responded to Sputnik I's launch by claiming he could have a satellite in orbit within 90 days of being given a go-ahead. His plan was to combine the existing Jupiter C rocket (confusingly, a Redstone adaptation, not a Jupiter) with the solid-fuel engines from the Vanguard, producing the Juno I. There was no immediate response while everyone waited for Vanguard to launch, but the continued delays in Vanguard and the November launch of Sputnik II resulted in the go-ahead being given that month. Von Braun kept his promise with the successful launch of Explorer I on 1 February 1958. Vanguard was finally successful on March 17, 1958.

ARPA selects Juno

Concerned that the Soviets continued to surprise the U.S. with technologies that seemed beyond their capabilities, the DoD studied the problem and concluded that it was primarily bureaucratic. As all of the branches of the military had their own research and development programs, there was considerable duplication and inter-service fighting for resources. Making matters worse, the DoD imposed its own Byzantine procurement and contracting rules, adding considerable overhead. To address these concerns, the DoD initiated the formation of a new research and development group focused on launch vehicles and given wide discretionary powers that cut across traditional Army/Navy/Air Force lines. The group was given the job of catching up to the Soviets in space technology as quickly as possible, using whatever technology it could, regardless of the origin. Formalized as Advanced Research Projects Agency (ARPA) on February 7, 1958, the group examined the DoD launcher requirements and compared the various approaches that were currently available.

At the same time that ABMA was drawing up the Super-Jupiter proposal, the Air Force was in the midst of working on their Titan C concept. The Air Force had gained valuable experience working with liquid hydrogen on the Lockheed CL-400 Suntan spy plane project and felt confident in their ability to use this volatile fuel for rockets. They had already accepted Krafft Ehricke's arguments that hydrogen was the only practical fuel for upper stages, and started the Centaur project based on the strength of these arguments. Titan C was a hydrogen-burning intermediate stage that would normally sit between the Titan lower and Centaur upper, or could be used without the Centaur for low-Earth orbit missiles like Dyna-Soar. However, as hydrogen is much less dense than "traditional" fuels then in use, especially kerosene, the upper stage would have to be fairly large in order to hold enough fuel. As the Atlas and Titan were both built at 120" diameters it would make sense to build Titan C at this diameter as well, but this would result in an unwieldy tall and skinny rocket with dubious strength and stability. Instead, Titan C proposed building the new stage at a larger 160" diameter, meaning it would be an entirely new rocket.

In comparison, the Super-Jupiter design was based on off-the-shelf components, with the exception of the E-1 engines. Although it too relied on the Centaur for high-altitude missions, the rocket was usable for low-Earth orbit without Centaur, which offered some flexibility in case Centaur ran into problems. ARPA agreed that the Juno proposal was more likely to meet the timeframes required, although they felt that there was no strong reason to use the E-1, and recommended a lower-risk approach here as well. ABMA responded with a new design, the Juno V (as a continuation of the Juno I and Juno II series of rockets, while Juno III and IV were unbuilt Atlas- and Titan-derived concepts), which replaced the four E-1 engines with eight H-1s, a much more modest upgrade of the existing S-3D already used on the Thor and Jupiter missiles, raising thrust from 150,000 to 188,000 lbf (670 to 840 kN). It was estimated that this approach would save as much as $60 million in development and cut as much as two years of R&D time.

Happy with the results of the redesign, on August 15, 1958, ARPA issued Order Number 14-59 that called on ABMA to:

Initiate a development program to provide a large space vehicle booster of approximately 1 500 000-lb. thrust based on a cluster of available rocket engines. The immediate goal of this program is to demonstrate a full-scale captive dynamic firing by the end of CY 1959.

This was followed on September 11, 1958, with another contract with Rocketdyne to start work on the H-1. On September 23, 1958, ARPA and the Army Ordnance Missile Command (AOMC) drew up an additional agreement enlarging the scope of the program, stating "In addition to the captive dynamic firing..., it is hereby agreed that this program should now be extended to provide for a propulsion flight test of this booster by approximately September 1960". Further, they wanted ABMA to produce three additional boosters, the last two of which would be "capable of placing limited payloads in orbit."

By this point, many in the ABMA group were already referring to the design as Saturn, as von Braun explained it as a reference to the planet after Jupiter. The name change became official in February 1959.

NASA involvement

In addition to ARPA, various groups within the US government had been considering the formation of a civilian agency to handle space exploration. After the Sputnik launch, these efforts gained urgency and were quickly moved forward. NASA was formed on July 29, 1958, and immediately set about studying the problem of crewed space flight, and the launchers needed to work in this field. One goal, even in this early stage, was a crewed lunar mission. At the time, the NASA panels felt that the direct ascent mission profile was the best approach; this placed a single very large spacecraft in orbit, which was capable of flying to the Moon, landing and returning to Earth. To launch such a large spacecraft, a new booster with much greater power would be needed; even the Saturn was not nearly large enough. NASA started examining a number of potential rocket designs under their Nova program.

NASA was not alone in studying crewed lunar missions. Von Braun had always expressed an interest in this goal, and had been studying what would be required for a lunar mission for some time. ABMA's Project Horizon proposed using fifteen Saturn launches to carry up spacecraft components and fuel that would be assembled in orbit to build a single very large lunar craft. This Earth orbit rendezvous mission profile required the least amount of booster capacity per launch, and was thus able to be carried out using the existing rocket design. This would be the first step towards a small crewed base on the moon, which would require several additional Saturn launches every month to supply it.

The Air Force had also started their Lunex Project in 1958, also with a goal of building a crewed lunar outpost. Like NASA, Lunex favored the direct ascent mode, and therefore required much larger boosters. As part of the project, they designed an entirely new rocket series known as the Space Launcher System, or SLS (not to be confused with the Space Launch System part of the Artemis program), which combined a number of solid-fuel boosters with either the Titan missile or a new custom booster stage to address a wide variety of launch weights. The smallest SLS vehicle consisted of a Titan and two strap-on solids, giving it performance similar to Titan C, allowing it to act as a launcher for Dyna-Soar. The largest used much larger solid-rockets and a much-enlarged booster for their direct ascent mission. Combinations in-between these extremes would be used for other satellite launching duties.

Silverstein Committee

Line drawings showing the evolution of the Saturn I rocket, from the original designs to the flown versions, and the uprated Saturn IB

A government commission, the "Saturn Vehicle Evaluation Committee" (better known as the Silverstein Committee), was assembled to recommend specific directions that NASA could take with the existing Army program. The committee recommended the development of new, hydrogen-burning upper stages for the Saturn, and outlined eight different configurations for heavy-lift boosters ranging from very low-risk solutions making heavy use of existing technology, to designs that relied on hardware that had not been developed yet, including the proposed new upper stage. The configurations were:

• Saturn A
  • A-1 – Saturn lower stage, Titan second stage, and Centaur third stage (von Braun's original concept).
  • A-2 – Saturn lower stage, proposed clustered Jupiter second stage, and Centaur third stage.
• Saturn B
  • B-1 – Saturn lower stage, proposed clustered Titan second stage, proposed S-IV third stage and Centaur fourth stage.
• Saturn C
  • C-1 – Saturn lower stage, proposed S-IV second stage (similar to the actual Saturn I).
  • C-2 – Saturn lower stage, proposed S-II second stage, proposed S-IV third stage.
  • C-3, C-4, and C-5 – all based on different variations of a new lower stage using F-1 engines, variations of proposed S-II second stages, and proposed S-IV third stages (with C-5 being similar to the actual Saturn V).

Contracts for the development of a new hydrogen-burning engine were given to Rocketdyne in 1960 and for the development of the Saturn IV stage to Douglas the same year.

Launch history

1965 graph showing cumulative history and projection of Saturn launches by month (along with Atlas and Titan)

Saturn Launch History 

PROGRAM	VEHICLE	MISSION	LAUNCH DATE	PAD
Saturn I	SA-1	SA-1	Oct 27, 1961	LC-34
Saturn I	SA-2	SA-2	Apr 25, 1962	34
Saturn I	SA-3	SA-3	Nov 16, 1962	34
Saturn I	SA-4	SA-4	Mar 28, 1963	34
Saturn I	SA-5	SA-5	Jan 29, 1964	LC-37B
Saturn I	SA-6	A-101	May 28, 1964	37B
Saturn I	SA-7	A-102	Sep 18, 1964	37B
Saturn I	SA-9	A-103	Feb 16, 1965	37B
Saturn I	SA-8	A-104	May 25, 1965	37B
Saturn I	SA-10	A-105	Jul 30, 1965	37B
Saturn IB	SA-201	AS-201	Feb 26, 1966	34
Saturn IB	SA-203	AS-203	Jul 5, 1966	37B
Saturn IB	SA-202	AS-202	Aug 25, 1966	34
Saturn V	SA-501	Apollo 4	Nov 9, 1967	LC-39A
Saturn IB	SA-204	Apollo 5	Jan 22, 1968	37B
Saturn V	SA-502	Apollo 6	Apr 4, 1968	39A
Saturn IB	SA-205	Apollo 7	Oct 11, 1968	34
Saturn V	SA-503	Apollo 8	Dec 21, 1968	39A
Saturn V	SA-504	Apollo 9	Mar 3, 1969	39A
Saturn V	SA-505	Apollo 10	May 18, 1969	LC-39B
Saturn V	SA-506	Apollo 11	Jul 16, 1969	39A
Saturn V	SA-507	Apollo 12	Nov 14, 1969	39A
Saturn V	SA-508	Apollo 13	Apr 11, 1970	39A
Saturn V	SA-509	Apollo 14	Jan 31, 1971	39A
Saturn V	SA-510	Apollo 15	Jul 26, 1971	39A
Saturn V	SA-511	Apollo 16	Apr 16, 1972	39A
Saturn V	SA-512	Apollo 17	Dec 7, 1972	39A
Saturn V	SA-513	Skylab 1	May 14, 1973	39A
Saturn IB	SA-206	Skylab 2	May 25, 1973	39B
Saturn IB	SA-207	Skylab 3	Jul 28, 1973	39B
Saturn IB	SA-208	Skylab 4	Nov 16, 1973	39B
Saturn IB	SA-210	ASTP	Jul 15, 1975	39B

Apollo program

Main article: Apollo program

The challenge that President John F. Kennedy put to NASA in May 1961 to put an astronaut on the Moon by the end of the decade put a sudden new urgency on the Saturn program. That year saw a flurry of activity as different means of reaching the Moon were evaluated.

Both the Nova and Saturn rockets, which shared a similar design and could share some parts, were evaluated for the mission. However, it was judged that the Saturn would be easier to get into production, since many of the components were designed to be air-transportable. Nova would require new factories for all the major stages, and there were serious concerns that they could not be completed in time. Saturn required only one new factory, for the largest of the proposed lower stages, and was selected primarily for that reason.

The Saturn C-5 (later given the name Saturn V), the most powerful of the Silverstein Committee's configurations, was selected as the most suitable design. At the time the mission mode had not been selected, so they chose the most powerful booster design in order to ensure that there would be ample power. Selection of the lunar orbit rendezvous method reduced the launch weight requirements below those of the Nova, into the C-5's range.

At this point, however, all three stages existed only on paper, and it was realized that it was very likely that the actual lunar spacecraft would be developed and ready for testing long before the booster. NASA, therefore, decided to also continue development of the C-1 (later Saturn I) as a test vehicle, since its lower stage was based on existing technology (Redstone and Jupiter tankage) and its upper stage was already in development. This would provide valuable testing for the S-IV as well as a launch platform for capsules and other components in low earth orbit.

The members of the Saturn family that were actually built were:

• Saturn I – ten rockets flew: five development flights, and five launches of boilerplate Apollo spacecraft and Pegasus micrometeoroid satellites.
• Saturn IB – nine launches; a refined version of the Saturn I with a more powerful first stage (designated the S-IB) and using the Saturn V's S-IVB as a second stage. These carried the first Apollo flight crew, plus three Skylab and one Apollo-Soyuz crews, into Earth orbit.
• Saturn V – 13 launches; the Moon rocket that sent Apollo astronauts to the Moon, and carried the Skylab space station into orbit.

• 

  A Saturn IB (AS-202) liftoff from LC-34

• 

  von Braun with the F-1 engines of the Saturn V first stage at the U.S. Space and Rocket Center

• 

  Rollout of Apollo 11's Saturn V to the launch pad