1.        ScopeThis article introduces the technology of launching satellites. 

2.        Background

In this article we generally refer to a man-made satellite as spacecraft prior to its delivery into orbit – except when the distinction is not obvious. For example – a spacecraft is mated with launch vehicle at the launch site; and a geostationary satellite is positioned at an altitude of 35786 km above the Earth’s equator.

The repetitive trajectory traversed by a celestial object around a central body (its parent) is known as its orbit.

For an object to attain a sustainable Earth orbit, the object must be propelled to an altitude above the Earth’s dense atmosphere and imparted a tangential velocity such that the object misses the Earth as it falls while remaining above the Earth’s dense atmosphere to avoid the intense frictional heat.

In absence of a tangential thrust (i.e., a thrust perpendicular to the Earth’s gravity), the object falls back to the Earth unless launched with sufficient thrust to break it free from the Earth’s gravity and continue its outward journey in space. Note the significance of tangential thrust – it must be imparted by the final stage of the launch vehicle (LV).

There are two general classes of orbit – open and close.

Open orbits are hyperbolic in shape. Satellites in such orbits escape the gravitational pull of the central body and continue to move in an ever-increasing distance from the parent. Open orbits are useful for inter-planetary exploration and travel.

Close orbits are elliptical wherein satellites continue to orbit the central body. Circular orbit is a special case of an elliptical orbit where eccentricity is zero. Close orbits are used invariably in all satellite applications – be it Earth observation, remote sensing, imaging, navigation, surveillance or communication.

Figure 1 is a representation of orbit classes around a central body illustrated as a circular dot. The shape of an orbit depends on the velocity V imparted to the satellite by the final stage of the LV at the release point P. The orbit acquires a circular shape when velocity V equals V_c, the velocity required to support a circular orbit at that altitude, as explained in the next section. When V is less than V_c the orbit takes an elliptical shape smaller than the circular orbit with the central body as the focus at the far end, i.e., P is at apogee; when V is more than V_c the ellipse is larger than the circle with the central body as the near focus, i.e., P is at perigee ; when V equals the escape velocity V_e (explained later), the orbit take shape of a parabola; for V more than the escape velocity orbits acquire  the shape of a hyperbola that gets flatter as velocity, V, is increased [1].  Note that the magnitude of velocity V imparted at the point P depends on the capability of the LV.

In this article we address close orbits. 

The velocity, v, required to achieve a circular orbit at an altitude R_s is given b the expression:

v = √ GM/r                       …. (1)

Where,

Gravitation constant, G = 6.67384 x 10^-11 Nm²/kg²

Mass of the Earth, M = 5.9723 x 10²⁴ kg

Satellite distance from geo-centre, r = R_E + R_s

Average equatorial Earth radius, R_E = 6378.15 km

Figure 2 illustrates the orbital velocity required for acquiring circular orbit at altitudes in the range 100-1000 km. For example, the velocity reduces from 7.25 km/s at an altitude of 200 km to 5.06 km/s at 1000 km. LVs must carry fuel to climb the additional 800 Km which nulls the advantage of a lower velocity at the higher altitude.

Escape velocity

Satellites escape the Earth’s gravitation when their kinetic energy exceeds the potential energy due to the Earth’s gravitation:

 ½ mv² > GMm/r                 …. (2)

m = mass of satellite, M = mass of the Earth, v = satellite velocity, r = satellite geo-centre distance

The escape velocity, v_e   is therefore,

 v_e > √ (GMm/r)                  …. (3)

For example, the escape velocity from the Earth at an orbital altitude of 200 km is > 11.18 km/s. Note its independence from satellite mass.

Preferred orbital altitudes

At low orbital altitudes atmospheric drag reduces orbital velocity and hence satellites lose altitude. The cumulative effect causes satellites to eventually enter dense atmosphere where a majority burn out due to frictional heat. Practical lower limit of sustainable orbits is ~180 km. A majority of operational satellites are positioned above 450 km where orbital decay due to drag has reduced significantly.

The Earth’s magnetic fields traps electrically charged particles around the Earth in doughnut shaped belts. These belts are known as Van Allen radiation belts after their discoverer. Radiation intensity in these belts varies gradually with altitude, increasing to the highest intensity between ~2,000–4,000 km followed by a reduction then rising again to high levels in between ~13,000–20,500 km before subsiding. High ion density in these regions affects electronics and on-board software/firmware adversely hence these regions are generally avoided for satellite systems.

Taking atmospheric drag as the lower limit and the first Van Allen radiation belt as the upper, the preferred orbital altitudes for low Earth orbit satellite systems lie within ~ 450-1,500 km. The next orbital windows lie between the first and second radiation belts (~ 10000-12000 km) and above the second Van Allen belt (~20500 km). Geostationary orbit at an altitude of 35786 km above the equator remains, by far, the most preferred choice for satellite communications. 

Other classes of orbits of interest in satellite systems include highly elliptical orbits and sun-synchronous orbit. Inclination of 63.4^o is of interest in serving high altitude regions due to its unique feature of maintaining the apogee fixed in space relative to the Earth. Using a repertoire of LVs, major launch service providers launch satellites to the entire genre.

3.       Launch Vehicle

The principle of rocket motion is based on Newton’s third law which states that to every motion there is an equal and opposite reaction. Hot gas obtained by burning a propellant in an ignition chamber ejected at a high velocity from a nozzle placed at the bottom of the engine propels the vehicle upwards in reaction.

A LV comprises an engine to provide thrust, a control and guidance system to keep the vehicle stable and on course, one or more payloads depending on the mission and structure to support all the subsystems.

Figure 3 shows the primary constituents of a LV. The structural systemcomprises an aerodynamically efficient streamlined enclosure of strong material like titanium or aluminium to support and keep all functional parts intact during launch. Some vehicles include fins at the bottom for aerodynamic stability. Due to the extremely high temperature generated in the rocket engine and high velocity acquired during launch (> 16,000 km/h) the vehicle gets extremely hot and hence heat resistant materials such as titanium and nickel alloys are used for casing.

The payload system comprises one or more mission dependent payload.

The guidance system provides the desired stability and control. It includes sensors and software to measure the orientation, location and velocity and apply corrections so that the vehicle remains steady and maintains the desired trajectory. Most modern rockets use gimbals that can rotate the nozzle to track the flight path and stabilise. Velocity control is applied by switching off thrusters when necessary. An inertial guidance system pinpoints the location and orientation of the rocket using readings from gyroscope and accelerometer by extrapolating initial location, velocity and orientation.

The propulsion system is dominant in launch functionalities and volume.  It comprises solid or liquid fuel or their combination, an oxidizer to provide oxygen to aid ignition, an ignition chamber to ignite the fuel and a nozzle at the bottom end of the chamber to discharge gas at a high velocity. Propellant performance is assessed by its specific impulse given as kg/kg/sec or, simply as sec. The performance efficiency of the rocket engine is determined by type and volume of the propellants taking into consideration the total weight of the vehicle including deadweight – parts that do not contribute to thrust. Payload to propellant mass including remaining part of the LV structure is known as payload ratio and should be as high as feasible [2]. In practice the ratio is of the order of 2.5-4.5 % – often less than 1%. For example, the payload ratio of Arian V and Soyuz 2 1-B are respectively 2.506% and 2.63% [3].

Thrust, F, produced by the engine depends on the exit velocity, pressure and mass of the gas flowing through the nozzle. The general thrust equation is given as

F = (dm/dt)_e*V_e – (dm/dt)₀*V₀ + (p_e– p₀)*A_e                                            …. (4)

Where  (dm/dt)_e = mass flow rate of gas at exit of nozzle

V_e = exit velocity of gas

(dm/dt)₀ = mass flow rate of a free stream of gas

V₀ = velocity of free stream

p_e = exit stream pressure

p₀ = free stream pressure

A_e = exit area of nozzle

The nozzle is designed to maximise mass flow rate, exit velocity, exit stream pressure and withstand extremely high temperatures.

Solid rocket engines are relatively simple, cost-effective and safe but difficult to control after ignition. Once started solid state motors cannot be stopped and continue to burn until depletion of the fuel. They are used to increase thrust during lift off; smaller motors are used in the final stage of launch vehicle; to lift satellites to higher orbit including geosynchronous orbit. The engine comprises a solid fuel combined with an oxidizer sealed along the side of a strong case with a perforation in the middle going down all the way to the nozzle. Since ignition takes place from the centre outwards towards the casing, the shape of the perforation determines the pattern and rate of ignition and consequently the generated thrust thereby a mechanism to control thrust.

Figure 4 illustrates a representative schematic (Adapted from [4]).

Solid propellants usually applied during launch belong to a family known as Composite – a heterogeneous mixture consisting of fuel, oxidizer and a binder sometimes including a catalyst to boost performance or compound to facilitate manufacturing.  

Solid motors have been used in Titan, Delta and space shuttle launch LV. Consider, as an example, the solid propellants used in the Space Shuttle [5]. The propellant mixture in each solid rocket booster motor consists of an oxidizer (ammonium perchlorate), fuel (aluminum), a catalyst (iron oxide), a polymer rubbery binder to hold the mix, and a curing agent. The mixture, known as ammonium perchlorate composite, is commonly used in boosters. Composed as a liquid mixture, it is cast into the desired shape and solidified into a rubbery solid. An 11-point star-shaped perforation is used in the forward motor segment and a double-truncated cone perforation is used in each of the aft segments and closure. This combination “provides high thrust at ignition and then reduces the thrust by approximately a third 50 seconds after lift-off to prevent overstressing the vehicle during maximum dynamic pressure.”[5]

Liquid rocket engines have the capability to control thrust. It comprises two tanks containing respectively cryogenic fuel and oxidizer. In a pump-fed system, each component is pumped into an ignition chamber through a command-controlled valve.  In a pressure-fed scheme high pressure gas is used to push the liquids into the chamber.  These super-cool constituents are circulated around the super-heated parts prior to entering chamber.

Figure 5 illustrates simplified schematic of a pump-fed liquid rocket engine.  

Liquid propellants are classified as petroleum, cryogenic and hypergolic [6].

Petroleum propellants are complex hydrocarbons refined from crude oil. Refined kerosene- known as RP-1 in USA – as fuel with liquid Oxygen at -183^oC as an oxidizer is used in many LVs such as in first stages of Atlas V, Delta II, Saturn 1B and Saturn V, all the stages of Falcon 1 and Falcon 9 and by several Chinese and Russian launchers

Cryogenic fuels use liquefied gases as fuel and oxidizer. Combination of liquid hydrogen at -253^oC as fuel and liquid oxygen as oxidizer is very efficient, although difficult to store for long periods and larger tank requirement of liquid hydrogen due to its low density, the propellant is used in many LVs due to its high efficiency. The propellant was used in the main engine of the space shuttle, and in upper stages of Saturn 1B, Saturn V. Centaur rockets, Atlas V, newer Delta IV rockets, Ariane rockets, etc.

 Hypergolic propellants have the property to ignite when fuel and oxidizer come on contact eliminating the need of ignition. Moreover since these propellants are easy to store, remaining liquid at normal temperature and readily controllable they are suited for spacecraft propulsion systems. Hydrazine and its derivatives such as Aerozine 50 with Nitrogen tetroxide or nitric acid are used in LVs such as Titan launch LV family and Delta II. [6]

Nitrogen tetroxide (N₂O₄) and hydrazine (N₂H₄) are useful in deep space rockets because of their long storability at achievable temperatures and pressures with simple ignition sequences. These propellants are toxic and hence require careful handling. Hydrogen peroxide, hydrazine, and nitrous oxide monopropellants – storable for long periods, simple to use, and suited for applications needing small impulses – are used for station-keeping; however, monopropellants provide lower specific impulse than bipropellants.

Ion-propulsion engines discharge high velocity positive ions instead of chemically generated gas to generate thrust, although the thrust generated is very small. These engines have a significantly higher thrust to propellant consumption ratio than chemical engines and are therefore widely used for low-thrust applications such as station keeping and propelling inter-planetary missions where space would be insufficient to carry enough chemical propellant. You learn how ion thrusters work here [Retrieved 15/09/2017].

The reader can view images of NASA’s F-1 rocket engine here. To date this engine remains the most powerful flown by NASA generating 1.5 million pound of thrust.  It was mounted at the bottom of Saturn V LV used in the Apollo programme.

Figure 6 shows the main features of Delta II LV (Courtesy NASA). Boeing’s Delta II is a reliable launcher recording nearly 100 consecutive successful launches [Delta II data sheet, Retrieved 19/09/2017].

Liquid oxygen and kerosene are in common use, e.g., in first stages of the Saturn V and Atlas V, all the stages of Falcon 1 and Falcon 9 and by several Chinese and Russian launchers

Liquid oxygen and liquid hydrogen are used in upper stage of Atlas V, and Saturn V, newer Delta IV rockets, Ariane rockets, etc.

Nitrogen tetroxide (N₂O₄) and hydrazine (N₂H₄) are useful in deep space rockets because of their long storability at achievable temperatures and pressures with simple ignition sequences. These propellants are toxic and hence require careful handling.

Hydrogen peroxide, hydrazine, and nitrous oxide monopropellants – storable for long periods, simple to use, and suited for applications needing small impulses – are used for station-keeping; however, monopropellants provide lower specific impulse than bipropellants [6].

Ion-propulsion engines discharge high velocity positive ions instead of chemically generated gas to generate thrust, although the thrust generated is very small. These engines have a significantly higher thrust to propellant consumption ratio than chemical engines and are therefore widely used for low-thrust applications such as station keeping and propelling inter-planetary missions where space would be insufficient to carry enough chemical propellant. You learn how ion thrusters work here (Retrieved 15/09/2017).

The reader can obtain a description and schematics of Atlas V LV here (Retrieved 31/08/2017).

4.      Staging

At lift-off LVs are fully loaded and hence at heaviest. An LV must therefore harness largest thrust at lift-off to carry its cargo further exacerbated by enormous atmospheric resistance encountered in the initial 100-150 km of ascent. Thrust at lift off can be of the order of 1.3 to 1.5 the weight of the vehicle. A single stage rocket is generally insufficient to achieve a successful launch. Multiple rocket stages, typically 2-3, can share the burden successfully by jettisoning the burnt stage to reduce the lift-mass of the next stage. The final velocity acquired by the payload is the sum of velocity increment contributed by each stage.

In series staging the succeeding stage is ignited on completion of the existing stage; in parallel staging smaller rockets are strapped to the side of the main rocket engine to operate in parallel. At burnout parallel boosters are ejected as the main rocket continues ascent. Hybrid staging uses both series and parallel stages.

In an equipotential field, the maximum velocity increment, Δv, that a LV of total mass, m₀, can impart is given as

 Δv =  v_g ln [1/(1-m_f/m₀)]                             …. (5)

Where v_g is the effective exhaust velocity of gas which depends on fuel and rocket nozzle design and m_f is the mass of expanded fuel.

Notice that the velocity increment increases by reducing LV mass, which occurs progressively with fuel depletion and ejection of spent parts of rocket.

5.       Improving launch efficiency

Depending on the LV capability and mission, launch cost to a LEO are of the order of US $3800/kg to US $20200/kg (year 2017 base) and to a GTO, ~ US $10500/kg to $27100/kg [7]. To reduce such high costs every possible effort is made to improve launch efficiency.

Enabling techniques include:

• Launch staging and jettisoning of used stage(s) – discussed in the preceding text;
• Judicious choice of propellants tailored to each stage;
• Optimum launch site selection;
• Piggy backing small SC;
• Multi-SC-multi-orbit launch;
• Air-borne Launch to reduce propellant during lift-off and provide launch location and weather independence;
• Use of ship-based launch pad to enable transport of the launch pad to a favourable location, e.g., near equator for an equatorial launch;
• Dedicated LVs for small SC:  According to a market survey at least 29 companies are actively involved in development of LVs for small SC (<500 kg) to service the renewed interest in small satellite applications, while in excess of 50 organisations are developing launch technology for SC <500 kg in mass [8].
• Reuse of jettisoned stages: Most LVs to date have used expendable rockets, where the jettisoned stages are lost – either burning in the atmosphere, although some parts may remain in space as debris, or are dropped in uninhabited areas like ocean. Parts of the jettisoned stages are expensive (tens of million) and must be rebuilt; if such components – rocket stages and fairings – were recovered and reused, a reduction in launch cost (e.g. 10-30%) would be feasible. A preliminary version of the technology has been introduced commercially in the USA by Space-X. Airbus has an ongoing program known as Adeline since 2010 (ADvanced Expendable Launcher with INnovative engine Economy) to promote a reusable launcher aiming to recover the main engines and avionics of the launcher, which represent 70-80% of the launch vehicle’s total value [25}. (Airbus is the largest aeronautics and space company in Europe). 

6.       Launch Vehicle Interfaces and capabilities

LVs incorporate a variety of mechanical and electrical interfaces to accommodate and interact with payload(s) during mission. Mechanical interfaces provide a secure connection of the payload(s) with its protective enclosure (known as, fairings) and mate the payload with the vehicle to facilitate stage separation and electrical connectivity. Electrical interfaces provide conduits to monitor SC and the LV during pre-launch preparations and post-launch, and assist in detecting stage separation and catastrophic failure. Ground controllers receive telemetered payload performance data and send commands from the ground via these interfaces.

Modern LVs are capable of launching auxiliary payloads in different orbits. For example, Atlas V launcher includes adapters for payloads ranging in weight from 10 kg to 500+ kg. In one instance its multi-picosatellite adapter was altered to launch 24 cubesats in the same mission.

The interested reader is encouraged to consult LV user handbooks to glean technical capabilities and program profile and if necessary define a preliminary mission profile including initial SC interfaces. For example, see Atlas [9] and Ariane [10] handbooks.

7.       Launch site

The minimum orbital inclination achievable depends on launch azimuth and latitude of launch station as follows:

 cos (i) = sin (ξ).cos (θ)                                    …. (6)

 Where, i is the minimum orbital inclination feasible; ξ, the launch azimuth and θ, the launch site latitude.

For an easterly launch (ξ = 90 deg.), the minimum inclination achievable equals latitude of the launch site, although in practice the inclination can be marginally higher.

Satellites require considerable amount of on-board fuel to alter their orbital inclination which increases satellite mass, size and more importantly the launch cost. Thus a launch site located on the equator or close to it is ideally suited to launch equatorial satellites and a northern site is better suited to launch satellites in inclined and polar orbits.

It is a practice to direct flight trajectory away from populated areas to eliminate the risk of damage due to LV debris in case of a launch failure or when a launch is aborted due to an anomaly. Launch sites are therefore often sited such that launch trajectories traverse over ocean. Examples are Kourou, French Guiana (5.1611° N) overseeing North Atlantic ocean to the east and Sriharikota, India (13.74⁰ N) adjacent to Bay of Bengal to the east, both suited for low inclination and equatorial launches; and Vandenberg Air Force station (34.742°) on west coast of US facing pacific ocean region to its west used for inclined orbit and polar orbit launches.

Figure 7 illustrates the sensitivity of minimum achievable inclination, i_m, to launch azimuth and latitude within 5^o– 60^o. Note that i_m:

• equals latitude of the launch site for an easterly launch;
• increases as azimuth is reduced, approaching a polar orbit for a northern launch in all cases;
• at lower latitudes always trails the greater for the same launch azimuth; however slope of i_m increases as latitude reduces.

An easterly launch benefits from the inertia of the Earth’s west to east rotation. The Earth moves at a speed, V_e, of about 1670 km/h (~464 m/s) at the equator (Note: The Earth rotates a distance of about 40070 km in 24 h).  The velocity reduces to about 328 m/s at latitude of 45^o and about 80.6 m/s at latitude of 80^o. 

Rotation velocity, V_l, at other latitudes is given as:

V_l = V_e Cos (θ_e) m/s                        …. (7)

Where, θ_e is the latitude. Figure 8 illustrates the Earth’s rotational speed at various latitudes. Notice that the figure is simply a cosine curve of amplitude V_e.

The predicted rotational speeds at a few representative launch facilities are as follows:

Kourou (ESA, French Guiana, 5.12⁰ N): 462 m/s;

Sriharikota (ISRO, India, 13.74⁰ N): 450.6 m/s;

Vandenberg (Air Force Base, USA, 34.74⁰ N): 381.2 m/s.

 7.1 Support ground segment

It is necessary to monitor the status of satellite throughout the launch. Hence support of several ground stations, located strategically for maximum visibility, becomes mandatory – particularly during the critical manoeuvres. This is often achieved by collaborative arrangements between space agencies and operators.

8.       Launch vehicle categories

The orbital altitude and payload carrying capability of an LV depends on its thrust generation capacity. Majority of LVs use expendable rockets where jettisoned stages of rockets are lost. Reusable launchers overcome this limitation by recovering and reusing one or several jettisoned parts such as the first stage tanks and fairings. Given that the main engines and avionics of the launcher represent 70-80% of the launch vehicle’s total value, reusability offers the potential of a sizeable reduction in launch cost

LVs are based on land, sea or air. Generally air-launch systems are suited to launch low mass satellites to LEO from any suitable airport. Sea-based LVs offer the option to launch from best possible location such as the equator for an equatorial orbit launch. Land-based launch sites are by far the most dominant offering numerous logistics and launch support advantages. However minimum achievable orbital inclination is fixed for each site.

9.       Launch vehicle selection criteria

Mission objectives set a guideline in terms of orbital parameters,  SC mass and dimension, availability, cost, vehicle capabilities, reliability, target trajectory, etc.[11]. Considerations include whether the spacecraft(s) would be a primary payload or piggybacked and whether the satellites form part of a constellation. The task is then to determine the most suitable vehicle that satisfies mission requirements optimally including geo-political factors and schedule compliance.

10. Launch campaign management

A launch campaign is managed and conducted by a dedicated team comprising experts of the SC and LV teams guiding the mission from the arrival of the SC on launch site through to lift-off and post launch analysis. The team manages schedule; ensures safe arrival and handling of the SC;  verifies compatibility of LV-SC interfaces; confirms that SC complies the stringent launch environment loads such as vibrations, acoustics and thermal; covers system engineering aspects including mission definition, trajectory optimisation, launch countdown, rocket staging, post-launch trajectory evaluation and compliance; and post-launch mission analysis.

Typically, a LV provider procures hardware from their industrial under a strict quality control regime and, when necessary, alters LV’s baseline design to comply mission requirements. The readiness of the assembled LV is assessed by the LV provider with participation of the SC team.

SC manufacturers/ operator consortium (the customer) and the LV provider liaise closely during the highly intricate launch process to ensure a faultless launch.

Consider Ariane 5 as an example [10]. Arianespace, a part of the European space transportation union, has capabilities to launch small, medium and large satellites in low, medium, highly elliptical geostationary transfer and earth escape orbits.

Typical mission tasks involve schedule consolidation, LV procurement and its adaptation for the mission, system engineering, launch campaign management with provision of safety and quality assurance to the customer at each stage. The completion lasts about 24 months typically.

The preparatory phase involves preparation of essential documentation and a technical verification to confirm that the given launch can be supported. The subsequent management and LV integration process is consolidated by a thorough and comprehensive documentation, developed and agreed in a series of pre-scheduled review meetings, where progress is monitored and corrections applied when necessary. LV adaptation is necessary when a launch requirement necessitates changes to the LV’s baseline configuration. For instance a multiple satellite launch requires additional adapters.

The launch campaign comprises activities beginning from the arrival of SC at the launch station through to countdown, launch, post-launch analysis and documentation. The activities include:

• Checking correct performance of the SC on arrival at launch site;
• Charging of the SC with propellants for on-station operations;
• Mating SC with the LV and checking readiness;
• Moving loaded LV to the launch pad – typically a day prior to lift-off;
• Check out of electrical and fluids umbilical to SC.

 Mission analysis

System engineering targets to ensure that the LV can complete the mission; the SC and interfaces are compatible with the LV; and LV customisation has been implemented. Further it addresses mission analysis, which involves: trajectory optimisation, orbit injection accuracy analysis; SC separation and collision avoidance analysis to minimise risk of collision between SC and LV’s housing module during separation, dynamic coupled loads analysis to ensure that SC is able to absorb various loads during launch. It ensures a harmonious radio environment by performing electromagnetic and RF compatibility analysis and thermal analysis to deal with exposure of the SC to thermal shocks.

Trajectory of a SC should be favourable relative to sun and visibility from telemetry, tracking and command (TT&C) stations during critical manoeuvres. For rendezvous operation it is necessary that the trajectory and launch time be chosen to converge to the designated target, e.g., International Space Station; or a designated location of the orbital arc, e.g., when the satellite forms a part of a constellation. Thus we observe that SC must be launched within preferred time slot known as launch window. Furthermore, we observe that the launch window is influenced by rocket’s characteristics [12]. More generally, launch window defines the most suitable time slot to launch taking into consideration all the mission constraints including safety.

Trajectories are tailored for each mission according to customer preference. For example, the customer may favour visibility of each launch event from supporting TT&C stations; or the customer may prefer the most fuel efficient trajectory. The complexity of trajectory and performance optimisation requires extensive modelling efforts involving orbital mechanics taking into considerations factors such as dynamic air pressure, drag coefficient and LV capabilities, while constraints include limitation of LV, SC, environment and location of the launch site. The optimisation criteria would depend on customer preference e.g., fuel efficiency and propellant margin. The analysis also forms the basis to assess telemetry coverage and RF link margins.

Guidance analysis confirms that the required guidance and navigation requirement of the mission is satisfied by the LV within tolerable performance limits. The analysis verifies that the guidance hardware and software can achieve acceptable performance under stringent conditions that include wide dispersion in LV trajectory,

Injection Accuracy Analysis provides estimates of orbit injection accuracy. Launch Window Analysis includes wind loading for ground and in-flight wind conditions. Flight safety analyses confirm safety in all respects. Electrical compatibility analysis aims to ensure thatelectrical circuit interfaces are compatible throughout. Post-flight data analysis consolidates the actual mission profile, obtained through telemetry and deduces deviations, against predictions. Destruct System Analysis evaluates robustness of flight termination system.

 11. Launch preparations

Preparations for the launch begin several days ahead of the day of lift-off (Day 0). Consider as an example, the activities counted down in days during the pre-launch phase of a dual launch mission of Ariane 5 [16]. Dual launch refers to launch of two SCs in the same mission.

Day        Task

-11

• Mating of upper SC to LV adapter

-10

• Transfer of mated SC to the final assembly area
• Integration of upper SC on LV’s payload housing (known as Sylda 5)
• Final checks and preparations of SC for launch

-9

• Mating of lower SC to LV adapter
• Transfer of mated lower SC to final assembly area

-8

• Installation of fairing on upper SC
• ntegration of lower SC on lower housing of LV

-7

• Functional checks and final preparations of lower SC and RF tests

-6  

• Integration of upper assembled parts on LV
• SC functional checks and RF tests.

-5

• Remaining SC functional checks and final RF link adjustments
• SC battery charging

-4  

• Dress rehearsal

-3

• LV final preparations
• SC battery recharging
• Fit and connect pyrotechnic devices to LV – known as arming (phase 1)
• Note: Pyrotechnic devices are used for separation (jettisoning) of used rocket parts and deployment of SC integral parts such as antenna. These single use devices shear a mechanical joint by a controlled electrically-generated tiny explosion on the device.

-2  

• Arming of LV (phase 2)
• Arming of the SC, when necessary
• Upper composite doors closure
• Note: Upper composite comprises (beginning from the bottom end, an LO₂-LH₂ liquid rocket engine followed by a vehicle equipment bay which supports guidance, stage sequencing, telemetry, tracking and safety systems.

 -1

• Transfer of LV to launch pad
• Launch chronology
• Filling of fuel in cryogenic main stage engine
• Filling of fuel in cryogenic upper stage engine
• Preparations for the final count-down begin about 9 hours prior to lift off
• The SC is checked out by conducting vital RF and functional tests to confirm readiness
• SC RF flight setup is configured with transmitter power set to specified level (H0-1h 30m) before lift-off (H0) and remains unchanged until 20 s after separation
• SC is switched on to internal power latest by H0-7 m
• automatic sequence starts nominally  at H0-7 m
• Countdown can be set to ‘hold’ in case of an anomaly and the countdown is then set back to H0-7 m
• SC can be switched back to external power in response if required
• SC stop action can be triggered up to H0-7 s

 In parallel, electrical checks are made to the LV concluding with filling of cryogenic propellants concluding at about H0-30 min.

 An example of countdown of space shuttle is available here  [13](Retrieved 12/07/2017).

The entire operation is regulated by a comprehensive safety regime to ensure safety in all respects of mission including flight safety following launch, prevention/managing accidents and concerns during hazardous operations such as propellant fuelling.

The mission, extending to LV hardware development and production including supplies from sub-contractors and component suppliers is compliant to internationally defined standards such as ISO 9001: V2000 to achieve the highest level of reliability and performance.

Furthermore LV providers exercise extreme care to minimise the impact of their activities on the Earth and space environments, extending to space debris management.

12.        Launch

We make a few general observations first.

• The separation of the booster rocket can be triggered by acceleration threshold detection.
• Fairing is released when drag due to molecular density of the atmosphere (aerothermal flux) drops below a threshold. A representative value is 1135 W/m2 for a geostationary transfer orbit (GTO) [10].
• After separation, the ejected part is oriented away from the orbit of the rocket to minimise risk of collision.
• The ejected part is spun to stabilize it for re-entry.
• Time to place a LEO satellite in orbit is of the order of 20 minutes (LV dependent).
• Time for place satellite in GTO is of the order of 30 minutes (LV dependent).
• A direct insertion Geosynchronous orbit (GSO) mission can place a satellite in GSO in about 5.75 hours (LV dependent).

12.1 LEO

Consider the mission sequence timeline of a sun-synchronous LEO launch from an Atlas V 401 [Minor adaptation of Table 2.4.1-1, [9]; Courtesy: United Launch Alliance]

Rocket engine (RD-180) Ignition	-2.7
RD-180 Engine Ready	0.0
Lift-off	1.10
Atlas Booster Engine Cut-off (BECO)	238
Atlas Booster/Centaur Separation	246
Centaur Main Engine Start 1 (MES1)	256
Payload fairing Jettison	264
Centaur Main Engine Cut-off (MECO1)	1015
Start Alignment to Separation Attitude	1017
Separate SC	1184
Start Turn to CCAM Attitude	1209
Centaur End of Mission	5584

Footnotes

1. Time is specified in seconds;
2. Fairing is released when drag has reduced to acceptable level  (measured as aerothermal flux e.g.,  < 1135 W/m2 for GTO);
3. Centaur is the upper stage liquid hydrogen- liquid oxygen engine of Atlas V launch system.

Table 1  Mission sequence timeline of a sun-synchronous LEO launch from an Atlas V 401 [Minor adaptation of Table 2.4.1-1, [9]; Courtesy: United Launch Alliance]

 The animation of launch sequence of a low Earth orbit satellite is available elsewhere on this site.

Air launch

SC is launched from an aircraft flying at a high altitude to minimise drag. For example, a US company (Orbital Science) uses an L-1011 Stargazer aircraft to launch satellites up to 977 lbs (443.16 kg) from an altitude of 39000 ft.  (11.89 km) to LEO [14].

12.2 Geostationary Transfer Orbit

The events below summarise the launch chronology to a GTO usingAtlas V 521 Standard GTO mission[Minor adaptation of Table 2.4.1-1, [9];

Courtesy: United Launch Alliance]
Rocket engine (designated RD-180) ignition	0:-2.71
Engine Ready (T = 0)	0:0
Solid Rocket Booster (SRB) Ignition	0:0.8
Lift-off	0:1.05
Strap-on SRB Jettison	1:58
Payload fairing Jettison	3:32
Atlas Booster Engine Cut off	4:09
Atlas Booster-Centaur Separation2	4:17
Centaur Main Engine Start 1 (MES1)	4:27
Centaur Main Engine Cutoff-1 (MECO1)	15:04
Start Turn to MES2 Attitude	17:12
MES2	23:22
MECO2	27:48.8
Start Alignment to Separation Attitude	27:50
Separate SC	30:37
Start Turn to CCAM3 Attitude	31:02
Centaur End of Mission	1:43:57

Footnotes

1)    Time format – H:M:S

2)    Centaur is the upper stage liquid hydrogen-liquid oxygen engine of Atlas V launch system.

3)    Collision and Contamination Avoidance Manoeuvre (CCAM) eliminates the risk of collision or contamination of satellite from Centaur after separation.

 Table 2  Launch chronology to a GTO usingAtlas V 521 Standard GTO mission [9].

12.3 Geostationary Earth orbit

Geostationary satellites can be launched directly to a Geostationary Earth orbit (GEO)or in stages via intermediate orbits.

Staged launch

In a staged launch  the satellite is initially placed in a low earth parking orbit followed by injection into an elliptical transfer orbit typically having an apogee at the altitude of GEO (35786 km), perigee at the parking orbit (e.g., 185 km for Atlas-Centaur) and an inclination of the parking orbit. A super-synchronous transfer orbit requires lower velocity increments to achieve GEO than the more traditional geosynchronous transfer orbit (e.g. , 48,789 km x 6545 km used for early GOES satellites – see bibliography)

GSO is acquired by imparting velocity increments in increments at the apogee aggregating the difference between the satellite velocity at the apogee and velocity required for a GSO. The aggregate velocity increment is minimal if it is applied at the point of least velocity i.e., apogee. Transportation between two coplanar orbits via an elliptical transfer orbit, known as Hohmann transfer after its inventor, requires the least velocity increment. The thrust for the transfer can be given by on-board thrusters known as apogee-kick motors. To acquire a GEO the inclination of the GSO must be reduced to near zero by imparting an appropriate velocity increment in a direction perpendicular to the orbital plane – usually applied at the same time as the circularisation manoeuvre.  The circularization and inclination reduction manoeuvres are often done in up to 3 separate manoeuvres to avoid long engine burns.

The satellite is made to drift in a near GEO orbit until it reaches the designated longitude where it is stopped by firing on-board thrusters to eliminate the residual inclination. When on station, satellites tend to drift gradually (over weeks) in longitude and inclination due to gravitational effects of the sun and the moon and perturbations caused by slightly oblate shape of the Earth.  Hence its location is adjusted regularly in manoeuvres known as East-West station keeping for longitudinal drift and North-South station keeping to correct inclination.

Figure 9 illustrates the launch sequence of a geostationary satellite pictorially. Notice the manoeuvres during the drift phase before the satellite reaches its designated longitude – antenna deployment, earth acquisition, solar array deployment and sun acquisition. Tests are carried out to ensure correct functioning of the satellite when stationed or at an intermediate location.

Figure 10 shows an example of the launch sequence of a geostationary satellite; the table below the figure lists typical time of major events [15].The precise launch flight plan depends on the launcher.

Table 3 illustrates the event time sequence of a direct insertion Geosynchronous orbit (GSO) mission for Atlas V High Load Vehicle (HLV) referenced to engine ready status (Minor adaptation of Table 2.4.1-1, [9]; Courtesy: United Launch Alliance)

Engine (RD-180) Ignition	-2.7
RD-180 Engine Ready (T = 0)	0.0
Lift-off	0.97
Liquid Rocket Booster (LRB) Engine Cut-off	3:48
LRB Jettison	3:56
Payload Fairing (PLF) Jettison	4:58
Atlas Booster Engine Cut-off (BECO)	6:07
Atlas Booster/Centaur Separation	6:15
Centaur Main Engine Start 1 (MES1)	6:25
Centaur Main Engine Cut-off (MECO1)	9:23
Start Turn to MES2 Attitude	13:38
MES2	19:48
MECO2	27:58
Start Turn to MES3 Attitude	5:32:37
MES3	5:38:47
MECO3	5:42:23
Start Alignment to Separation Attitude	5:42:25
Separate SC	5:45:12
Start Turn to CCAM1 Attitude	5:45:37
Centaur End of Mission	6:58:32

 

Footnote

Collision and Contamination Avoidance Manoeuvre (CCAM) eliminates the risk of collision or contamination of SC from Centaur after satellite separation.

 Table 3  Event time sequence of a direct insertion GSO mission [9].

12.4 Large constellation

Deployment of large non-GEO constellations comprising dozens to hundreds of satellites requires a series of multiple-satellite launches.

Here we introduce deployment of two first generation constellations – Iridium and Globalstar. [Note: The first generation satellites are in midst of replacement by second generation satellites.]

Deployment of non-GEO constellations within a short span economically is vital for commercial success. Therefore multiple satellite launches in quick succession is necessary. Satellites can be launched in low or medium earth orbit directly or in two stages via an intermediate parking orbit. For the latter the launcher deposits clusters of satellites at regular intervals in a parking orbit. When a precise orbital position has been determined, each satellite is moved to the desired orbit by imparting the desired thrust vector at a specified time by tele-commanding on-board thrusters. Support of a network of tracking stations dispersed strategically across the world is necessary for satellite monitoring and orbit determination.

Iridium first generation constellation comprised 66 satellites in six 780 Km altitude circular polar orbital plane and Globalstar first generation constellation comprised 56 (included 8 in-orbit spares) satellites in eight 1414 km altitude inclined orbital planes. Both consortiums targeted the deployment of their respective constellations in 15-18 months.

Consider Delta II launch of Iridium first generation satellites from Vandenberg Airforce Base in California, USA [16, 17]. Satellites were launched in groups of three. The first satellite was jettisoned at an altitude of 638 Km in 3130 seconds (52 min 10 sec) after lift-off, followed by releasing other two in succession at intervals of 200 seconds. The satellites gradually drifted apart over several days. About an hour and forty minutes after launch the first radio contact with each satellite was made by controllers stationed at Motorola’s satellite control centre in Chandler, Arizona when satellites were in view of Iridium’s tracking station in Oahu, Hawaii to check whether satellites were in the designated orbit and functioning normally. There were four tracking stations located in Hawaii, Yellowknife and Iqaluit in Northwest territory of Canada, and Snjoholt in Iceland. To avoid the possibility of exhausting on-board batteries solar arrays were deployed at an early stage, followed by deployment of communication antennas for ground and inter-satellite communications. About three hours in to the mission, secondary antennas were tested in an initial checkout. The satellites switched from the secondary antenna to the main antennas in the fifth orbit giving an increased communications throughput. In the first two days satellite’s primary antennas and modems were checked out, batteries recharged, software upgraded if required, and feeder link and inter-satellite link performance checked out. About 48 hours in to the mission, satellites began ascending to their respective final location at an altitude of 780 Km by firing low-thrust electro-thermal hydrazine thrusters. Thrusters were fired firstly over poles to raise the orbit and then over equator to circularise the orbit. The final orbit was acquired gradually in about two weeks.

Globalstar planned their deployment through a number of different launchers. The first two launches in batches of four were to be launched from Cape Canaveral Air Force Station in Florida, each on a single Delta 2 LV. The next three launches were planned in batches of 12 on Zenit LV. The final three launches were to be launched on Soyuz LV, supplied by a Russian-French consortium, ferrying 4 satellites per launch. Some readjustments were made due to a launch failure. Note that failure of a multi-satellite launch has a severe impact on mission schedule and cost. One Zenit launch failure resulted in a loss of 25% of Globalstar constellation.

For Delta 2 launch, four satellites held in a canister in the LV were jettisoned almost simultaneously at an altitude of 1250 Km. The first four Globalstar satellites were placed in the same orbital plane to simplify the launch. The initial few manoeuvres were initiated by each satellite autonomously using on board computers.  The manoeuvres included extension of the magnetometer boom, acquisition of sun and the Earth, stabilisation to avoid tumbling, and deployment of solar arrays to avoid battery depletion. When in orbit, the satellites were controlled by Globalstar’s Operations Control Centre in San Jose, California using tracking stations in Texas, France, South Korea and Australia. A preliminary health checkout of each satellite was made to ensure that the vital functions such as attitude control and propulsion behave normally. Within a few hours each satellite was commanded to fire thrusters to jettison itself to its final altitude of 1400 Km. Satellite injected last was boosted first to minimise the risk of collision.

Zenit vehicle launches twelve satellites simultaneously delivering them to an altitude of 920 Km. Satellites were held in canisters as with Delta launch, and ejected within 4 seconds in rapid succession. Speed is essential to minimise the risk of placing satellites in an incorrect orbit due to movement of the canister. As only six satellites must be placed in each plane, the satellites were grouped and injected in three separate planes. The initial manoeuvres were identical to the Delta launch. But as satellites were at a lower orbit they travelled faster giving only 10-12 minutes of visibility from ground stations. A further consideration was that the satellites underwent different radiation and thermal condition at 920 km than experienced at 1400 km for which they were optimised. Hence the satellites were moved to the higher altitude as soon as possible. Satellite altitude was altered in groups of two or three to minimise the work load on ground controllers. Satellites were allowed to orbit until they reached their respective orbital plane.

Soyuz launch also jettisons satellites to the same altitude as Zenit but only four satellites were launched at a time. The eight satellites launched from Soyuz launches were intended to remain partially activated as spares in 920 km orbit.

12.5 Small satellites

There has been a surge of interest by academic, government, scientific, and commercial organisations throughout the world in utilizing the vastly increased capabilities of tiny, low-cost satellites realized by advances in miniaturisation of space electronics and sensors.

Table 4 [18] specifies categorization of satellites by weight, which although not standardised as such, is widely used by the small satellite community.  

Micro	10-100 kg
Nano	1-10 kg
Pico	0.1-1 kg
Femto	<0.1 kg

Table 4  Categorization of tiny satellites by weight [18].

As an aside, cubesats – used by students, amateur and professional organisations for missions ranging from participation incentive for students to technology evaluation, scientific experiments, earth observation, imaging, etc. – have gained considerable world-wide attention. A cubesat is a 10 cm cube nano satellite of mass up to 1.33 kg – its architecture standardised to facilitate cost reduction and interface with LV deployment mechanisms [19].

A hindrance in the uptake of tiny satellite missions remains the launch cost which can be far greater than the cost of satellite itself and budget of many such missions. The present launch options of tiny satellites are – dedicated, rideshare and piggybacked.  

Dedicated launch is an option for deployment of large constellations in a single or multiple orbital plane or for satellites targeting a particular orbit and schedule. Being a relatively expensive option, it is used by government and commercial organisations with well-defined mission goals, targeted technology and little fund constraint. Satellites are launched singly or in groups as discussed in the example of the preceding section.

In a rideshare approach, payloads from different sources share a single launch to reduce their respective launch cost. All the payloads must be compatible with LV’s mechanical and electrical interfaces. The choice of orbits is constrained by the capability of the LV.

In a piggybacked launch the target satellite(s) is a secondary payload sharing the mission with a primary payload. Sharing reduces cost but limits orbit and schedule flexibility since the launch is driven by mission goals of the primary satellite. Numerous constrains on the size and design of piggybacked satellites are applied to protect the primary payload. The interface to the LV gnaws into satellites mass, size and cost budget; this could, for example, adversely impact the capacity of satellites’ solar panel; or limit the volume and capability of the propulsion system. The choice of orbit would be restrained by and dependent on the primary payload mission goals. Mandatory insurance when imposed further aggravates the financial overhead.  

Various organisations and companies have evolved their LV designs to launch tiny satellites at low cost including a provision to launch in multiple orbits. For example, in February 2017, an Indian Polar Satellite LV (PSLV-C37) launched a 714 kg Cartosat-2 series remote sensing satellite along with 103 Nano satellites of different sources averaging about 6.5 kg/satellite. The aggregate mass of the payloads was 1378 kg [20].

Nevertheless shortcomings and high demand of existing launch systems for numerous other missions necessitates dedicated and affordable launch facilities for tiny satellites – an observation corroborated by market analysis [21]. One estimate predicts up to 3000 nano and microsatellite (1-50 kg mass) launch between 2016 and 2022 at an average growth rate of 13% per year [22]. There is thus considerable interest in providing launch services to this segment with cost reduction of several orders of magnitude [8].

Technologies to lower launch cost and provide rapid and economic turnaround, while eliminating existing constraints include [24]:

• Mass production of rocket parts;
• Improved nozzle design, known as airspike, to provide  uniform performance at all altitudes, where traditional bell nozzle fall short [23];
• Improved monopropellant that offers better reliability, improved performance, more compact design at lower cost;
• Air launch to provide flexibility in selection of launch site to suit inclination, lower weather impact, eliminate elaborate launch site preparations and reduce propellants;
• Manufacture and deployment of customized cubesats in space within the International Space Station, using 3D printer and ferried non-printable components thereby eliminating need of earth launch, simplify satellite design and minimise deployment delays;
• Flexible software catering for multiple LVs;
• Reusable space tugs stationed in space to transport satellites from low earth orbit to their final destination precluding the need of on-board propulsion to make satellite “lighter, less complex and cheaper to manufacture” [25].

Many LVs launch SC to a single orbital plane. One of the challenges in deploying multi-plane constellation of pico-nano sized satellite is the difficulty in distributing satellites to their respective orbital plane due to the limited propellant budget of such satellites coupled with a limited monetary budget that prohibit multiple launches. To distribute them in different orbital plane following a single launch requires significant on-board fuel, which increases the mass, size and cost of satellites to unacceptable levels. Similarly, launching a cluster separately in each plane increases the deployment cost prohibitively.

Innovative solutions include utilization of nodal precession of orbit caused by the Earth’s perturbations to spread out orbital planes and achieving plane separation via Lissajous orbit around Earth-Moon Lagrange point L1 [26].

Dependence of nodal precession, Ω, of the orbit on satellite to geo-centre distance r, semi-major axis a, inclination i, and eccentricity e, is approximated as:

Ω= 9.95 [r/a]^3.5[ cos (i) / ( 1-e²)² ] deg./day                           …. (8)

Following a multiple-satellite single plane launch a, i, e orbital parameters of each satellite is adjusted in a sequential time-separated sequence such that each satellite is deposited in the desired orbital plane. The scheme can be further optimised by encapsulating satellites in orbital groups and moving each group into their designated plane where satellites alter orbital parameters individually or collectively with support from the mother craft. Drag on each satellite can be adjusted by deploying frictional surfaces to achieve same-plane orbital separation. However, the time taken to deploy the constellation by the nodal precession method could run into months.

Lissajous orbits are quasi-periodic which revolve around the Langragian point L₁, a point on Earth-moon line where gravitational forces from them are equal. Satellites require only a nominal amount of fuel to retain this orbit. Satellites of a multi-orbit constellation are placed in such an orbit in a single launch with satellite groups of each orbit encapsulated separately. Next, satellites belonging to each orbital plane are propelled back to their respective designated orbital plane one at a time. The final orbit is attained by slowing down each satellite by a manoeuvre known as aerocapture where drag resistance is applied by deploying friction-inducing surface; followed by a circularisation manoeuvre executed by satellite’s propulsion system. This method is said to require less fuel than each satellite altering orbital plane following a single coplanar launch [26].

13. Non-rocket launch concepts

Numerous non-conventional launch concepts have been proposed over the years. Presently such concepts are early developmental stages [26]. The argument in favour of pursuing such concepts is that perhaps some of them can be utilized in tandem with conventional techniques to reduce launch costs. An easily affordable access to space unravels the possibilities of numerous space-based endeavours such as space colonies, space-based power generation, space travel, space manufacturing, and a myriad scientific and technological endeavours.

LIST OF ABBREVIATIONS

BECO	Booster Engine Cut-off
CCAM	Collision and Contamination Avoidance Manoeuvre
GEO	Geostationary Earth Orbit
GSO	Geosynchronous Earth Orbit
GTO	Geostationary Transfer Orbit
LEO	Low Earth Orbit
LRB	Liquid Rocket Booster
LV	Launch Vehicle
MECO1	Main Engine Cut-off 1
MECO2	Main Engine Cut-off 2
MEO	Medium Earth Orbit
MES1	Main Engine Start 1
MES2	Main Engine Start 2
SC	Satellite
SRB	Solid Rocket Booster
TT&C	Telemetery Tracking and Command

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