In 1957, the Soviet Union placed the first man-made object in orbit around the earth. Since then, numerous launch vehicles have been developed to improve the performance, reliability, and cost of placing objects in orbit. By one estimate, roughly 75 active space launch vehicles either have established flight records or are planning an inaugural launch within the year. This does not include the numerous launch vehicles from around the world that are no longer operational such as the Jupiter, Redstone, Juno, Saturn, Scout, Thor, Vanguard, and Cones-toga family of rockets from the United States or the N-1 from the former Soviet Union, to name just a few. Despite the many differences among all of these launch vehicles from both past and present, one common element can be found in all but four of them: they are ground-launched. Of the four exceptions, two are air-launched (NOTSNIK and Pegasus), one is ship-launched (Sea Launch), and one is submarine-launched (Shtil). It is important to keep in mind that numerous air-launched and ship-launched suborbital launch systems are in use by militaries, commercial entities, and educational institutions. However, the four mentioned are the only mobile launch systems that can place objects into a sustainable Earth orbit.

Mobile Space-Launched Vehicles

Project Pilot (NOTSNIK). NOTSNIK is the oldest and, until recently, the least well known of the four mobile space-launched systems. Following the launch of Sputnik by the Soviet Union, President Eisenhower’s administration elicited proposals to launch a satellite into orbit. The Naval Ordinance Test
Station (NOTS) located at China Lake in California proposed launching a rocket from a jet fighter (1). The idea is the same as that of the current Pegasus vehicle: reduce the amount of energy needed to place a payload into orbit by launching it above the denser portion of the atmosphere. In this fashion, the engineers at NOTS designed a vehicle from existing rocket motors that could place a 2-pound satellite in a 1500-mile-high orbit. The engineers recognized the energy savings from such a launch concept and also the utility of such a flexible platform. Launching from a jet fighter could, theoretically, place a satellite into any orbit from anywhere in the world at any time.
The U.S. Navy accepted the proposal from NOTS in 1958, by some accounts as a safety net in the event that the ongoing Vanguard project was unsuccessful. The program was officially called Project Pilot, but the engineers at NOTS preferred the name NOTSNIK in direct reference to the Soviet satellite that was currently orbiting above them and the rest of the world. A Douglas Aircraft F4D-1 Skyray was the carrier aircraft for the rocket and consequently was considered the first stage. The second and third stages were modified antisubmarine missiles. The final stage was taken from a Vanguard rocket. The entire launch vehicle measured a mere 14 feet in length and had four fins at the aft end that provided a span of 5 feet.
The NOTSNIK was launched six times from an altitude of about 41,000 ft. Four of those launches ended in known failures. However, the results of two have never been verified. Some in the program insist that they achieved their goal of placing the small payload of diagnostic instruments in orbit. At least one ground station in New Zealand picked up a signal in the right place at the right time. However, confirmation that the signal was from the NOTSNIK payload was never established. Even the possibility of a success was veiled in secrecy for more than 40 years for, by all accounts, two critical reasons. The first was that in the days following the early embarrassments of Vanguard, the Eisenhower administration did not want to claim success unless it was absolutely certain. The second reason was that a mobile air-launched system that could reach orbit had extremely appealing military applications. However, the tactical advantages of such a system were far outweighed by the strategic consequences, as stated in the Antiballistic Missile (ABM) Treaty between the United States and the former Soviet Union that was concluded in 1972 (2):
Further, to decrease the pressures of technological change and its unsettling impact on the strategic balance, both sides agree to prohibit development, testing, or deployment of sea-based, air-based, or space-based ABM systems and their components, along with mobile land-based ABM systems. Should future technology bring forth new ABM systems ‘based on other physical principles’ than those employed in current systems, it was agreed that limiting such systems would be discussed, in accordance with the Treaty’s provisions for consultation and amendment.
Pegasus. Roughly 30 years later, while NOTSNIK remained an official government secret, the idea of launching payloads into space from an airborne platform was revisited in the form of the Pegasus launch vehicle. The driving forces behind NOTSNIK and Pegasus were essentially the same. An air-launched space vehicle provides several advantages compared with ground-based counterparts.
As an example, Pegasus is launched at an altitude of 39,000 ft, which is above a significant portion of the atmosphere. As mentioned, with NOTSNIK, this eliminates the need for extra performance that would otherwise be needed to overcome atmospheric forces. This also implies that the structural components of the vehicle can be lighter, which improves the efficiency of the rocket as a whole. The energy required from the launch vehicle is also reduced by the speed already achieved by the carrier aircraft. An air-launched system also allows applying more of the impulse of the first stage along the velocity vector. This is a more efficient use of the vehicle’s energy than that of ground-launched vehicles that must first apply the thrust almost perpendicular to the velocity vector already imparted by Earth’s rotation. These factors combine to produce a requirement for a velocity increment that is on the order of 10% less than a comparable ground-launched rocket.
The Pegasus vehicle is a winged, three-stage, solid rocket booster (Fig. 1). It is the first space-launched vehicle developed solely with commercial funding. Three versions have been developed and flown over the years: Standard, Hybrid, and XL. The XL is the only vehicle within the Pegasus family currently in production. The XL is roughly 10,000 lbm heavier than the Standard or Hybrid models and is roughly 6 ft longer. Because the XL extends farther aft beneath the L-1011 carrier aircraft, the port and starboard fins become an obstacle to the landing gear doors. To correct this problem, the port and starboard fins were modified to include an anhedral of 23°. To maintain commonality between the various members of the Pegasus family of vehicles, the same anhedral was introduced into the Standard vehicle, which was then given the designation Pegasus Hybrid. Other than the anhedral of the fins, the Standard and Hybrid vehicles are exactly the same. The Standard, the first Pegasus vehicle built, was flown on six missions. The Hybrid vehicle has flown four times. The XL vehicle has flown 21 times. Of 31 Pegasus launches, only three missions failed to reach orbit.
Figure 1. Disassembled version of standard Pegasus launch vehicle.
Figure 1. Disassembled version of standard Pegasus launch vehicle.
The Pegasus XL was designed and developed to provide increased performance above and beyond that provided by the Standard and Hybrid vehicles. A typical Pegasus XL vehicle weighs roughly 51,000 lbm at launch, is 55.4 ft long and 50 inches in diameter, and the wingspan is 22 ft (3). At launch, the Pegasus XL is carried aloft by the company’s carrier aircraft, a modified L-1011, which originally saw commercial service with Air Canada. The vehicle is dropped from an altitude of 39,000 ft at Mach 0.8. Five seconds after release from the L-1011, the first stage ignites and the vehicle’s on-board flight computer continues the sequence of events that eventually lead to orbital insertion. The brief coast period between drop and stage one ignition is designed to provide a safe distance between the L-1011 and the launch vehicle.
The Pegasus Standard vehicle was originally dropped from a NASA-owned and operated B-52. The Pegasus vehicle was attached to one of the pylons underneath the starboard wing much in the same manner as the early supersonic and hypersonic test vehicles such as the X-15. For a variety of reasons, Orbital purchased and modified the L-1011 to facilitate all future launches.
Unlike the B-52 that supported initial Pegasus launches, the L-1011 carries the Pegasus vehicle underneath the fuselage rather than underneath the wing. Once Pegasus is ready to be mated to the carrier aircraft, it is towed from Or-bital’s integration facility at VAFB to the plane on the Assembly and Integration Trailer (AIT). Regardless of where the launch is to take place, the Pegasus is always integrated and mated to the L-1011 at VAFB. From there, the launch system can travel to any location in the world for launch. There is enough ground clearance for the L-1011 to take off and land with Pegasus attached underneath. However, the added height of the AIT underneath Pegasus requires raising the L-1011 off the ground slightly by hydraulic jacks to mate Pegasus to the carrier aircraft (Fig. 2). While mated to the L-1011, the vertical rudder actually protrudes into the plane’s fuselage in a compartment specifically designed for this purpose. When mating the Pegasus to the L-1011, the rudder is usually detached from the Pegasus vehicle and placed inside the housing first. Then the Pegasus is rolled underneath the L-1011 and attached to the rudder and then to the plane. Removing the rudder first minimizes the height to which the L-1011 needs to be raised for the mating process. The entire mating process from rollout to mating takes about 6 hours. Pegasus is attached to the L-1011 using four hooks on the center box of the wing and a fifth hook on the forward portion of the vehicle. The inside of the airplane has been stripped of all unnecessary equipment and hardware. Up front in what would normally be the first class cabin are eight seats for personnel during ferry flights from VAFB to the launch site of interest and two computer stations from which personnel can monitor the health of the vehicle and the payload. The rest of the interior of the cabin has been completely gutted. Access to the rear portion of the aircraft cabin is obtained through a galley door.
Figure 2. Fully assembled Pegasus launch vehicle being mated to the L-1011 aircraft. This figure is available in full color at
Figure 2. Fully assembled Pegasus launch vehicle being mated to the L-1011 aircraft.
Unlike most other launch vehicles in the U.S. fleet, the Pegasus launch vehicle is integrated horizontally on the AIT (Fig. 3). Horizontal integration facilitates easy access to the vehicle and eliminates the need for high bays and large cranes. Components are received as needed either from groups within Orbital Sciences or from outside vendors. To ensure that all of the major flight hardware and software is thoroughly tested before flight, Pegasus, like many other vehicles, is subjected to a series of ”fly to orbit” simulations at various stages of the integration process. Four flight tests are normally performed. The first tests the three stages individually. The second test is conducted after the three stages are electrically mated together. The third test is performed after the three stages are electrically and mechanically mated and the stack is electrically mated to the payload. The fourth and final flight test is performed once the payload has been mechanically mated to the rest of the vehicle and the half of the fairing that includes the pyro devices necessary for jettisoning the shroud is electrically mated. These tests are intended to verify that various systems function and also respond as expected to known disturbances. If the inertial measurement unit (IMU) onboard receives data to indicate that an unexpected attitude change has occurred, will the fins or thrust vector control systems respond accordingly? Are all the commands to the various subsystems appropriate, and do those subsystems respond appropriately? Once the Pegasus vehicle has been mated to the L-1011 carrier aircraft, one last test is performed, called the Combined Systems Test (CST). This test verifies that the launch vehicle and the carrier aircraft are communicating as expected. This is particularly important since the vehicle’s health can be monitored both from telemetry that is broadcast from the vehicle to the ground via antennas on Pegasus and also by the computer stations inside the L-1011 via hardwired electrical connections. More importantly, some data and commands are sent to the Pegasus vehicle before launch. The only method currently available for accomplishing this transfer of data is through the electrical connections between the Pegasus vehicle and the carrier aircraft.
Horizontal integration of Pegasus launch vehicle. This figure is available in full color at
Figure 3. Horizontal integration of Pegasus launch vehicle.
To be fully mobile, the Pegasus launch system must also be fully self-contained. Except for those services provided by the range (such as radar coverage), the L-1011 can transport all of the equipment required to support a launch of Pegasus, including, of course, Pegasus itself (Fig. 4). Some launches take place off the coast of California where the Western Range (based at VAFB) is the lead range. In these instances, no ferry flight is required. The L-1011 simply takes off from VAFB and flies to the designated drop point roughly 100 nmi out to sea. The checklist that is processed in the control room on the day of launch requires about 4 to 5 hours to complete. The L-1011 usually takes off an hour before the scheduled launch time. If all systems are ”go,” as determined by the mission team members in the control room, the launch conductor on the ground commands the pilot of the L-1011 to drop the Pegasus from the carrier aircraft. Shtil. In a classic example of turning swords into plowshares, the Russian Navy developed a satellite delivery system for nonmilitary applications that uses a submarine-launched. The SS-N-23 (NATO’s designation) is a three-stage liquid-fueled vehicle that can deliver small satellites to low Earth orbit. Very little is known about this launch vehicle service including performance to various altitudes and inclinations. What is known is that two satellites belonging to the Technical University of Berlin were successfully launched in 1998 from a Russian submarine for the stunningly low price of $150,000 (4). Some sources indicate that the typical commercial price for a Shtil launch is actually in the neighborhood of $500,000. There are two possible reasons for the low cost of a Shtil launch. The first is that more than 200 missiles have already been produced by the Russian military. There is also speculation that offering commercial launch services provides a way to maintain proficiency in launching missiles without using precious military funding. One disadvantage of this system is that the Shitl vehicle likely does not have enough performance to achieve circular orbits in the medium to high Low Earth Orbit (LEO) altitudes (4). This is a direct result of the Shtil’s heritage as a ballistic missile first and foremost. Sea Launch. The most recent mobile launch system is the Sea Launch vehicle which is launched from a converted oil-drilling platform along the equator (Fig. 5). Sea Launch is both the name of the launch vehicle and the name of the international joint venture that provides the launch services. The partnership is comprised of Boeing, KB Yuzhnoye of Ukraine, which provides the two Zenit stages, and RSC Energia of Russia, which provides the Block DM-SL upper stage. The launch vehicle and payload integration takes place at the vehicle’s home port of Long Beach, California. Once integration is complete, the launch vehicle is loaded onto the converted oil-drilling platform and towed to a predetermined launch location at the equator, specifically 154° West. Once on site, the Zenit 3SL is raised into its launch attitude (vertical) and launched. A second ship that houses mission personnel and the control room monitors the launch from nearby. The vehicle itself is a little less than 200 ft long and roughly 13 ft in diameter. The performance to Geosynchronous Transfer Orbit (GTO) is approximately 5250 kg (4). ”In terms of spacecraft mass in final orbit, this would be equivalent to approximately 6000 kg of payload capability if launched from Cape Canaveral, because the spacecraft does not need to perform a plane change maneuver during the Geosynchronous Earth Orbit (GEO) circularization burn” (5).
L-1011 aircraft taking off with Pegasus. This figure is available in full color at
Figure 4. L-1011 aircraft taking off with Pegasus.

Computer simulation of Sea Launch. This figure is available in full color at
Figure 5. Computer simulation of Sea Launch.
There are three key phases in the integration of a Sea Launch vehicle (5). Phase I takes place in the Payload Processing Facility (PPF). This phase includes receipt of the spacecraft, processing of the spacecraft, testing, and enclosure within the payload fairing. Phase II takes place on the Assembly and Command Ship (ACS). This entails mating the encapsulated spacecraft to the launch vehicle and testing the integrated stack. Phase III takes place on the Launch Platform (LP) once the vehicle has been transferred from the ACS. While still in port, the integrated launch vehicle is raised to its vertical launch attitude so that a series of tests can be conducted. The launch vehicle is then lowered back into a horizontal position, stored in an environmentally controlled room, and transported to the equator while on board the launch platform. At the launch site, the launch vehicle is rolled out to the launch pad, raised to a vertical attitude again, and fueled. The launch is performed by an automated system and monitored by the Assembly and Command Ship which is moved for launch to a distance 6.5 km away (Fig. 6).
The Assembly and Command Ship for Sea Launch serves as the launch vehicle integration and testing facility. In addition to acting as the temporary home for launch crews, the ship also houses the Launch Control Center (LCC) and the equipment necessary to track the initial ascent of the rocket. Unlike the
Figure 6. Sea Launch successfully lifts DIRECTV 1-R satellite into orbit. This figure is available in full color at
Figure 6. Sea Launch successfully lifts DIRECTV 1-R satellite into orbit.
Pegasus carrier aircraft that was modified after serving in a different capacity, the ACS was designed and constructed specifically to suit the unique requirements of Sea Launch. The ship is roughly 660 ft long and 110 ft in beam and has an overall displacement of approximately 30,830 tonnes.
The rocket assembly facility is on the main deck of the ACS where the launch vehicle integration takes place. This activity is conducted before setting sail for the equator and simultaneously with spacecraft processing. After the spacecraft has been satisfactorily processed, it is encapsulated and transferred to the rocket assembly compartment, where it is mated to the launch vehicle. Following integration and preliminary testing, the integrated launch vehicle is transferred to the launch platform. Then both ships begin the journey to the equator, which takes roughly 12 days.
The launch platform has all of the necessary systems for positioning and fueling the launch vehicle, as well as conducting the launch operations. Once the launch vehicle has been erected and all tests are complete, personnel are evacuated from the launch platform to the ACS using a link bridge between the vessels or a helicopter. Redundant radio-frequency links between the vessels permit personnel on the ACS to control all aspects of the launch, even when the command ship has retreated to a safe distance before launch. The launch platform, which was converted from an oil drilling platform, is very stable. It is supported by a pair of large pontoons and is propelled by a four-screw propulsion system (two in each aft lower hull). Once at the launch location, the pontoons are submerged to a depth of 70.5 ft to achieve a more stable attitude for launch, level to within approximately 1°.

Advantages of Mobile Space-Launched Systems

NOTSNIK, Pegasus, Sea Launch, and Shtil were never intended to replace the existing fleet of ground-launched rockets. Rather, they effectively supplement the existing worldwide capability by providing additional services to a targeted market of payloads that benefit greatly from the mobility and flexibility of these unique space-launch systems. These vehicles can provide services similar to ground-launched vehicles for payloads within their weightclass. In fact, all four vehicles have fixed launch locations for standard services. For example, Pegasus uses the launch location of 36° N, 237° E for all high-inclination missions that originate from VAFB. In this regard, the mobile launch systems are no different from ground-launched vehicles in that they repeatedly launch from a fixed location, albeit a location that is not on land. However, they can also offer services and performance that avoid many of the restrictions inherent in being constrained to a particular launch site. Few of those restrictions are trivial. They include inclination restrictions, large plane changes required to achieve low-inclination orbits from high-latitude launch sites, large plane changes required to transfer from GTO to GEO when launching from certain ranges, and low-frequency launch opportunities for missions that require phasing such as those involving a rendezvous with another spacecraft already in orbit. Inclination Restrictions. Inclination restrictions stem from range safety considerations. To understand these restrictions fully, it is first necessary to understand two concepts: (1) transfer orbits and (2) instantaneous impact-point tracks.
Transfer Orbits. Transfer orbits are intermediate orbits established by the various stages of a launch vehicle that provide a path to the final desired orbit. The transfer orbits for early stages are mostly suborbital, meaning that some portion of the orbit intersects Earth’s surface. The most efficient way to transfer between two orbits is to apply thrust at opposite apses. An application of thrust in the right direction at the perigee of the initial orbit will raise the apogee. Coasting to the new apogee and applying thrust (again in the appropriate direction) at this apsis will then raise the perigee. This provides a stair-step approach to raising the altitude of a vehicle’s orbit. The ascent of a launch vehicle from launch to orbit follows a similar trend with one critical caveat. The impulse of initial stages is usually not sufficient, individually, to raise the perigee above Earth’s surface. This means that using the optimal Hohmann transfer approach would bring the launch vehicle back to Earth before another transfer burn could be made. As a result, initial launch vehicle stages usually apply their thrust at places within a transfer orbit other than the apses and usually always on the ascending side of the orbit.
Consider a modest three stage, ground-launched rocket launching into a circular low Earth orbit as an example. Before launch, the vehicle is effectively sitting at the apogee of an orbit (Fig. 7). If the surface of Earth were not present to support the rocket, it would be drawn downward along a path that would take it closer and closer to Earth’s center before swinging back to an apogee altitude equal to the radius of Earth. This is essentially the first of several transfer orbits and the rocket has not even been launched. When the rocket lifts off, it applies its thrust at an apsis, but in a direction that is perpendicular to the initial velocity vector of the rocket, which itself is in the direction of Earth’s rotation. During the first burn, the vehicle slowly tilts over so that the thrust is applied in a direction that is increasingly parallel to Earth (Fig. 8). This has the effect of increasing both the apogee and perigee. The perigee will most likely still be suborbital at the end of the burn. The apogee will be increased sufficiently that the launch vehicle can coast up to a location near the new apogee, following the first stage burnout, and ignite the second stage. The key consideration here is that the second stage will be ignited near but not at the apogee. Again, this is not the most energy-efficient way to transfer orbits, but it is necessary because the opposite apsis is still below Earth’s surface, and the second stage may not have sufficient impulse to raise it above the atmosphere. Igniting the second stage at a location other than the apogee again has the effect of raising both the perigee and the apogee. In this case, because only one stage is left, the burn is designed to raise the apogee to the desired altitude of the final orbit. After the second stage burns out, the vehicle coasts up to the new apogee and ignites the third stage. This will raise the perigee up to the final orbit altitude without changing the altitude of the apogee.
Figure 7. Path of rocket without Earth's surface. This figure is available in full color at
Figure 7. Path of rocket without Earth’s surface.
 Path of rocket after launch. This figure is available in full color at http:// www. mrw. interscience.wiley. com/esst.
Figure 8. Path of rocket after launch.
Impact-Point Tracks. By always burning on the ascending side of the trajectory and iteratively raising the apogee while the transfer orbit remains suborbital, anything jettisoned before the final burn will reenter the atmosphere and either burn up or impact Earth’s surface. As the burn of each stage progresses, the point at which the transfer orbit intersects the Earth extends farther and further downrange until, at some point late in the final burn, there is no longer a point of intersection. These points of intersection comprise the instantaneous impact-point track. Clearly, as the vehicle is coasting, the instantaneous impact point does not change. Conversely, during a motor burn, it is constantly changing and each point represents the location of impact on Earth if, in fact, the thrust were to be instantly terminated either by design or due to some sort of failure. It is this impact-point track and the need for it to avoid populated areas that is a primary source of inclination restrictions from various ranges.
For any rocket launch, whether it be space-based, suborbital, ground-launched, ship-launched, or air-launched, the public-safety considerations that must be satisfied are very stringent. Those stages of a rocket that are jettisoned before reaching orbit should avoid land. And no launch vehicle whose impact-point track nominally crosses land can risk a casualty among the public with a probability of greater than 30 in a million. Calculating the expectation of a casualty depends on many factors, including the reliability of the launch vehicle (e.g., how many failures it has had in the past), the density of the population being over flown, and the speed with which the instantaneous impact-point track crosses over a populated region. Late in flight, the distance between successive impact points increases dramatically and reduces the risk to the population below. This is why it is generally more permissible to overfly populated regions far downrange than it is early in flight. For instance, the risk to a populated region in Africa from a rocket launched at the Eastern Range would, in general, be less than the risk posed to an area with the same population density over flown in the Caribbean. This is not to say that over flight of any part of Africa is acceptable. There are some extremely high population densities in Africa, especially along the west coast of northern Africa, which are avoided at all costs. And it is this very consideration that constrains the paths of many launch vehicles from the existing ranges.
The key land masses that must be avoided early in flight for vehicles launching from the Eastern Range include the entire eastern seaboard of the United States when launching on an ascending pass (northerly direction) and the Caribbean and South America when launching on a descending pass (southerly direction). For maximum performance from any given launch vehicle, this restricts the range of inclinations achievable from the Eastern Range to between roughly 28.5° and 51° for ascending passes and between 28.5° and 40° for descending passes. Clearly, inclinations outside this range would be achievable if plane changes were instituted, but that has the disadvantage of reducing the maximum available performance for any given launch vehicle. Higher inclinations are available from the Western Range but restrictions still exist there due to Hawaii, islands in the South Pacific, and the western coasts ofboth North and South America.
When the inclinations from both the Eastern and Western Ranges are combined (assuming direct injection), a block of inclinations is unavailable without plane changes and subsequent reductions in weight-to-orbit capabilities. For small payloads with limited budgets that require an inclination outside what is directly available from the existing ranges, the cost of launching on the heavy-lift launchers that can execute the necessary plane changes can be prohibitive. And reducing launch costs by flying as a secondary or even tertiary payload is advantageous only in the rare event that a primary payload can be found that requires the same final orbit. For these customers, Pegasus and Shtil provide an alternative due to their relatively low cost, mobility, and self-contained launch infrastructure. Sea Launch provides a similar alternative for the heaviest satellites that are intended for either GEO or low Earth orbits. Plane Changes Required to Achieve Low Inclinations. The inclination of an orbit represents the angle between the equatorial plane and the orbital plane around Earth. This also happens to be similar to the definition of lines of latitude. It is no coincidence then that the maximum latitude ofthe ground track for any object in space is roughly equivalent to the inclination of the object’s orbit. The only reason that the maximum latitude is not exactly equal to the inclination is because Earth is not a perfect sphere. Conversely, this implies that the minimum inclination attainable by a launch vehicle is roughly equivalent to the latitude of the location from which it is launched. The maximum is 180° minus the latitude of the launch point. This leads to the important conclusion that the only latitude from which all inclinations are directly accessible is 0° (the equator). The Eastern Range is at a latitude of roughly 28.5°. Therefore, the minimum inclination attainable without plane changes is roughly 28.5°. Lower inclinations can be achieved by launching into any available inclination, achieving a preliminary orbit, and then making an inclination correction burn when the satellite is over the equator or at any latitude that is numerically less than the desired inclination. The significant disadvantage of this process is that inclination changes while in orbit require a great deal of energy. The larger the change in inclination required, the more energy must be expended. Depending on the final orbit desired, this usually requires an additional stage to correct the inclination and achieve the final orbit. The most common recipient of this type of orbit maneuver is a satellite headed to geosynchronous orbit. However, there are low Earth orbit payloads that require low inclinations as well. The ability of Pegasus and Shtil to move the drop point to a latitude from which such energy-intensive plane changes would not be required permits smaller launch vehicles to achieve the same orbit from lower latitudes that larger vehicles can achieve from higher latitudes. The difference in cost, complexity, and performance can often mean the difference for some customers between launching or not.
Some launch locations maintained by other countries are at significantly lower latitudes than those in the United States. For some customers, such ranges can provide the necessary services. However, many satellites in the United States, especially government sponsored, are required to contract with a U.S. launch service provider and use a U.S. controlled range.
Phasing. An object’s orbit is essentially a locus of points that defines the path of the satellite. Those points define a plane that goes through the center of Earth. To define an object’s precise position within an orbit, that plane and every position in it is defined with respect to both Earth and a coordinate system, one of whose axes always points toward the vernal equinox. Every position of a satellite as it orbits Earth is defined in terms of an epoch (time), the semi major axis, and eccentricity, measured from Earth’s center, inclination and argument of perigee, which are both referenced to Earth’s equator, and the right ascension of the ascending node, which is referenced to the vernal equinox frame.
A rendezvous between two objects in space involves a series of maneuvers designed to make the orbital elements of both objects the same, hence confirming the fact that they have, in fact, become a single object orbiting Earth. Just as motor burns can raise or lower the perigee or apogee of an orbit or change the inclination, so too can motor burns be used to change every orbital element that defines a satellite’s motion. However, changing some of those elements, especially those that require plane changes, requires large amounts of energy, and they are considered ”expensive” in the parlance of orbital mechanics. One way to avoid paying the high price of actively changing the orbit of a satellite with a motor burn is to do it passively through the aid of various external forces. Several naturally occurring forces cause every orbital element to change over time. These include atmospheric drag, solar radiative pressure, the gravitational attraction of the Moon, Sun, and planets, and the nonuniform gravitational forces due to Earth’s oblateness. These forces can be used to one’s advantage when planning a rendezvous mission. However, some changes resulting from these forces can take a very long time to reach significant levels. This means that the initial differences between the rendezvousing satellite and the target must be initially small to avoid spending too much unproductive time in orbit. This can be accomplished by simply timing the launch appropriately so that at the time of orbital insertion, the satellite that has newly arrived in orbit is very close to the orbital plane of the target satellite.
To accomplish this maneuver, the launch must occur when the target satellite passes almost directly overhead. It also must be passing in the same direction as the intended launch. In other words, if the satellite being launched is to head off in a southerly direction (along the descending pass), the target satellite must be overhead and also on its descending pass as well. Otherwise, the two satellites will end up with right ascensions that are 180° apart which would be excessively expensive (either in terms of time or energy) to correct once in orbit.
For ground-launched vehicles, the wait between successive passes of the target satellite could be as much as several days, depending on the target orbit because the distance between ground tracks on successive passes depends on the period of the orbit, which depends on the orbit’s semimajor axis. Clearly, the ground track of an object that requires only 90 minutes to orbit Earth will be more closely spaced than the ground track of an object whose period is several hours. These ground tracks will pass to the east and west of the given launch site on a daily basis, but the distance between the ground track and the launch site will only be minimized by a periodicity of the order of days.
Mobile assets, however, can eliminate the wait by essentially choosing a launch point that is ideally suited for a rendezvous. Instead of waiting for the ground track to come to the launch point, the launch point is moved to the ground track. In this way, the launch opportunities can be reduced from one every two to three days to at least once a day if not twice a day, if the launch vehicles have the flexibility to launch on both ascending and descending passes.
Consider an example of a satellite being launched by a Pegasus XL to rendezvous with a satellite currently in orbit at an altitude of 400 km circular. A normal ground-launched vehicle would require a wait of about 2 days between successive launch attempts. However, the mobility of Pegasus permits two launch opportunities every day, which is graphically represented in Figs. 9 and 10. Two key assumptions need to be kept in mind when viewing these figures. The first is that the maximum range of the Pegasus carrier aircraft is roughly 1000 nmi. This includes a captive carry to the launch site, an aborted launch, and a return to base with Pegasus still attached. The second assumption is that for launches that do not require the full advantage of Pegasus’ mobility, the standard launch point for Pegasus out of the Eastern Range is 28° N, 281.5° E. The vertical axes in Figs. 9 and 10 represent the difference in argument of latitude between the two satellites (the angular separation within the same orbital plane). The horizontal axis represents the launch point as the difference in degrees from the nominal point listed before. The diagonal lines represent the difference in argument of latitude for each day in the first week of October, which was chosen simply as an example. Figure 9 represents the difference in argument of latitude for northerly launches (launch along the ascending pass). Figure 10 represents the difference in argument of latitude for southerly launches (launch along the descending pass). The horizontal lines simply demarcate zero angular separation between the two satellites.
Figure 9. Graph of difference in argument of latitude for northerly launches.
Figure 9. Graph of difference in argument of latitude for northerly launches.
The intersection of a diagonal line with a horizontal line defines a drop point within the range of the Pegasus carrier aircraft from which Pegasus can be launched and effectively deliver its satellite to the front door of the target satellite at the time of orbital insertion. Realistically, this is not how a rendezvous would normally be achieved. Ideally, the satellite being launched would be placed in a temporary parking orbit slightly below and behind the target satellite. Over the course of several orbits the distance separating the two objects would be slowly decreased using several controlled burns of the satellite just placed in orbit. This would imply that a drop point is needed not to achieve 0° difference in argument of latitude but some finite value. The example is still valid. Simply shift the horizontal lines up or down until the desired difference in argument of latitude is matched. Again, an intersection between a diagonal line and the horizontal line defines a launch point within the range of the Pegasus carrier aircraft that would result in the desired difference in argument of latitude.
Graph of difference in argument of latitude for southerly launches.
Figure 10. Graph of difference in argument of latitude for southerly launches.
As can be seen from Figs. 9 and 10, every day except two in the first week of October provides two launch opportunities. A southerly launch on the 3rd does not provide a drop point within the range of the carrier aircraft that will achieve the desired result. However, a drop point can be found on that day if the launch is along the ascending pass instead. Likewise, a suitable drop point cannot be found within the range of the carrier aircraft on the 6th of October when launching along the ascending pass, but one can be found if launching in a southerly direction. The same qualitative results would be obtained for any other time frame. The quantitative results might be slightly different. For instance, instead of having only one launch opportunity on the 3rd and 6th it may be the 4th and the 7th. But the end result is the same. The mobility of Pegasus and, by definition, Sea Launch, and Shtil provides ideal rendezvous launch opportunities at least once a day and in most cases twice a day.
Clearly there are disadvantages with all of these mobile assets. Pegasus is limited in its size due to the restrictions ofthe L-1011 and, more importantly, the mechanical limitations of the hooks that hold the vehicle to the plane. Sea Launch has somewhat of a temporal disadvantage in that it requires almost 2 weeks to travel to the launch site. Those problems are exacerbated for Shtil because its home port is farther north. Nonetheless, for some specific missions, the mobility and flexibility that are provided by these unique space-launched assets provide valuable supplemental services to the fleet of existing ground-launched vehicles.

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