Preliminary design for a small-lift launch vehicle pylon with Air launch mechanism.

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Dipartimento di Ingegneria Civile e Industriale

Corso di laurea Magistrale in Ingegneria Aerospaziale

Tesi di laurea

MASTER’S DEGREE COURSE IN AEROSPACE ENGINEERING

Preliminary design for a small-lift launch

vehicle pylon with Air launch mechanism.

19 February 2019 Anno Accademico 2018-2019

Candidato:

Anuraj Kodakkottil

Relatore:

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Abstract

The Brobdingnagian increase in the global space Industry as the result of the entry by multiple private giant corporates into the commercial space domain and the massive technological advancements in the field of small and nanosatellites opens up a market where the small satellites no longer need to be carried as piggybacks and with compromised launch timings. Air launch mechanisms are expected to be the key players in this sector considering the quantum of work demonstrated in recent years. The thesis work tries to address the issue of designing and optimization of the multiple options available for air launch mechanism concentrating on using existing aeroplanes modified to act as a carrier for the launchers. A compilation of theoretical works available on this technology and modern engineering tools are used to develop a preliminary design of the pylon for air launch concept. A conceptual design of the pylon integrating the launcher and aircraft is designed using classical methods, different Structural, Aerodynamic, Modal analysis are done to validate the design and are iterated for better performance. A basic fracture mechanics estimation is also done both theoretically and using simulation tools. The thesis aims at providing a path for the technical aspects for air launch mechanisms serving both academic and commercial purposes and deduces a preliminary design for the pylon connecting the launch vehicle and carrier aircraft to analyze different mechanical properties.

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Acknowledgment

I would like to express my deepest appreciation to everybody who provided me the possibility to complete this work.

Prof Mario Rosario Chiarelli, my supervisor and mentor for this thesis for providing his knowledge, guidance and making time promptly without any condition, without whom this project would have never been completed. The alma mater University of Pisa, the academic and non-academic staffs at my University for providing adequate resources whenever its needed. My colleague Varghese Jacob, who will be doing the future works and extension of this thesis for the continuous assistance and support throughout the process.

The people near my bosom, Helen, Joshi, Lan, Lineesh and Titto for always being ready to go the extra mile for me. All my friends who made the campus a home away from home, especially Dimitri, Noemi, Rafael, Beatriz and Avantika for proving me that friendships know no language, Sahar, Arjun, Sanjo, Rahul, Albie, Blessen, Shreepali, Sai, Kesha, Simha, Karthik, Arun and to everybody unnamed who have spread their happiness along the way.

Last, but not the least my Mom, Dad and my family, their love, support, sacrifice and encouragement throughout my education to whom I dedicate this work and the rest of my life.

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CHAPTER 1 - INTRODUCTION

1.1 Global Space Industry

The global space market in the past few years has been showing the trend of engaging more into commercial and corporate domain compared to the monopoly of government research and defense activities earlier which were the main investors in the market. The globalization and liberalization have led for the opening of a window of opportunity for smaller industries and organizations in the space industry.

The size of the global space industry, which includes satellite services, ground equipment, government space budgets, and global navigation satellite services (GNSS) equipment, is estimated to be about $335 billion [1]._{The satellite television represents the largest segment of}

the activity with $98 billion in revenues which is about 29 per cent of the total budget. The next biggest Industry following satellite television are services enabled by global navigation satellite systems (GNSS), which is the size of about $81 billion in revenues. The global government space budgets represent $77 billion, or 23 per cent of the total. Other satellite services (fixed and mobile satellite services, broadband, and remote sensing) generated about $30 billion in revenues, and ground equipment represents $28 billion in revenues. Satellite manufacturing generated nearly $17 billion. All this activity would not be possible without orbital launch services. Global launch services are estimated to account for $5.4 billion of the $335 billion total [1]_{. Most of this launch activity is captive; that is, the majority of payload}

operators had existing agreements with launch service providers. Since the emergence of small and Nanosatellite technology the scenario has changed drastically, creating a window for small scale industries and launch providers to innovate and attract major Angel investors.

Similar to ORBCOMM and Iridium which led to the commercial perspective of a need for small launchers in the 1980s and 90s resulting in the Pegasus development, cubesats and new constellations such as oneweb, Planet and small startups from multiple developing countries like India, China, Malaysia etc. Has resulted in the demand for the same now. As small satellite capability increases, the traditional ride shares and secondary payload opportunities available to operators can no longer satisfy the dynamic market with the time and positional requirements. During the past decade, there has been an increasing trend of interest in having new, lower cost, dedicated small launchers especially triggering multiple corporate giants like Virgin group, Space-X etc. To invest more on developing launchers specially dedicated to catering the needs of small-sats.

The thesis work tries to address the issue of providing a low cost, flexible, easily handled launch mechanism for Small and Nanosatellites. A variety of options available in the academic and Industrial domain is considered, studied, compared and air-launch method is selected as a result to design a conceptual solution for the structural integration on a readily available aircraft in the market. Multiple aspects of the financial and commercial issues are also addressed, and an estimate of the budget and cost is generated. A major portion of the thesis work is concentrated on the design, analysis and validation of the structure.

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1.2 Market study

The Small satellite market as discussed earlier is already on the limelight for multiple reasons, even though an exact value for the total revenue is not available due to the policies of multiple developers. A compilation of the available data up to 2018 is analyzed and are discussed in detail below. 2018 Nano-Microsatellite Market Forecast by Space Works Commercial is taken as the base of the market survey which is also backed by the report of Small Launch Vehicles – A 2018 State of the Industry Survey by Northrop Grumman Corporation [1]_.

Spaceworks Commercial in their 2018 Nano-Microsatellite Market Forecast projects up to 2,600 nano-microsatellites launching in the upcoming 5 years. This growth comprises of the growth in private investment dollars and government interest throughout the world, but especially in the United States. There is also a perceived shortage of launch opportunities with many of the new entrants habitually quoting a “2-year backlog” on existing vehicles as a potential differentiator for their own endeavor. Even a small portion of this market can be a game changer for most of the industries

entering this field.

Most of the already ‘involved in space programs’ countries are contributing to this market by assigning different programs especially with this aim in mind.

The United States has been one of the first few countries to show interest in this field, The U.S. Department of Defense (DOD) and NASA have noticeably increased the attention into the small launcher market. As small satellites increase in utility, efficiency and capability, dod and its associated agencies are interested mostly in “launch on demand” services along with traditional launch services, but also. Programs such as DARPA’s Airborn Launch Assist Space Access (ALASA) and NASA’s Venture Class Launch Services (VCLS) promised to fund new entrants in their development of small launch vehicles. The thesis works have

inherited multiple results from these studies considering the quantity of the innovation and research output these programs has contributed can’t be ignored.

European governments have their share of contribution too as ESA’s Future Launchers Preparatory Programme (FLPP) and studies funded through the European Union’s Horizon 2020 have both contributed needed investment in the European market [2]_{. Individual countries}

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such as the United Kingdom has also taken a new interest in small satellites by actively pursuing potential launch sites for many of the new entrants.

China also is in the forefront of this market, even though the political situation currently prevailing in the US doesn’t allow any US operators to use Chinese launchers, the rest of the world is still utilizing them.

Organization Vehicle Name Country First Launch

China Aerospace Science and Technology Corporation

Chang Zheng 11 China 25 Sep 2015

Rocket Lab Electron USA/New

Zealand

21 Jan 2018

China Aerospace Science and Technology Corporation

Kaituozhe-2 China 3 Mar 2017

Ex-Pace Kuaizhou-1A China 9 Jan 2017

Northrop Grumman Minotaur I USA 27 Jan 2000

Northrop Grumman Pegasus XL USA 5 April 1990

Table 1. 1 Operational Small Launch Vehicles [1]

1.3 State of the art

A number of other Vehicles were under development during the time as the thesis was under processes, some of them which follow the Air launch method and a similar configuration. Launcher One [20]_{, which is the latest addition to this group by Virgin Inc. Is a very interesting} project which has successfully tested the pylon structure while the report is generated, the after effect of this project can guide the future works of this project in an efficacious manner. A synthesized collection of the Launch vehicles under development is provided on Table 1.2 a number of organizations have not updated their estimated date for the first launch, and this date now lies in the past.

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Organization Vehicle Name Country Latest

Launch Date

ABL Space Systems RS1 USA Q3 2020

Aphelion Orbitals Helios USA 2021

Bagaveev Corporation Bagaveev USA 2019

Bspace Volant USA 2018

Celestia Aerospace Sagitarius Space Arrow CM

Spain 2016

Cloud IX Unknown USA

CONAE Tronador II Argentina 2020

Cubecab Cab-3A USA 2020

Departamento de Ciencia e Tecnologia Aeroespacial

VLM-1 Brazil 2019

ESA Space Rider Europe 2020

Firefly Aerospace Firefly Alpha USA Q3 2019

Gilmour Space

Technologies

Eris Australia/Singapore Q4 2020

Interorbital Systems NEPTUNE N1 USA

ISRO PSLV Light India Q1 2019

Land Space Landspace-1 China H2 2018

Launcher Rocket-1 USA 2025

LEO Launcher Chariot USA Q4 2018

Link Space Aerospace Technology Group

Newline-1 China 2020

One Space Technology OS-M1 China 2018

Orbex Orbex United Kingdom

Orbital Access Orbital 500R United Kingdom 2020

PLD Space Arion 2 Spain 3Q 2021

Rocket crafters Intrepid-1 USA Q1 2019

Rocket Star Star-Lord USA 2018

Sky rora Skyrora XL UK/Ukraine

Space Ops Rocky 1 Australia 2019

Space LS Prometheus-1 United Kingdom Q4 2017

Spin Launch Unknown USA

Stofiel Aerospace Boreas-Hermes USA 2019

Strato Launch Pegasus (Strato) USA

Vector Space Systems Vector-R USA H2 2018

Virgin Orbit Launcherone USA H1 2018

Zero2infinity Bloostar Spain 2017

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1.4 Investments

As discussed earlier, Investment is another key area which resulted in the sudden growth of the small satellite industries. The age-old tradition of only the Government and Publicly funded organizations has cleared the way for an inclusive private funded project culture. Different successful startups such as spacex [21]_{has been a motivation for multiple entrepreneurs to step} into the field of Aerospace. A different dominant organization having wide experience in Automotive, Mechanical and similar Industries are also stepping in aiming at commercial success. All the identified external sources of funding for each vehicle are shown in Table 1.3. Due to strategic reasons, few companies keep this information under tight control, and thus it is not available on the public platform.

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Organization Funding Source

Aphelion Orbitals Angel investors

Bagaveev Corporation Tim Draper, Adam Draper, DCVC, New Gen Silicon Valley Partners, Wei Guo, Data Collective, Sand Hill Angels

Celestia Aerospace One signed up customer

Cube Cab Biz Plan Competition, Self funded

ESA ESA

Expace 8 investment institutions Gilmour Space

Technologies

Blackbird Ventures, 500 Startups

Interorbital Systems Self, Presales

Land Space Angel Investors; Series B (all from non-government)

One Space Technology Legend Holdings, HIT Robot Group at Harbin Institute of Technology, Chun Xiao Capital, Land Stone Capital

Orbex High-Tech Gründerfonds, private investors, the UK Space Agency and the European Commission Horizon 2020 programme

PLD Space Spanish government, EC, Caixa Capital Risc, Gobierno de Aragon, GMV, ESA, Gonzalo de la Pena, EC

Rocket Lab NZ Gov, Kholsa, VBP, K1W1, LM, Promus Ventures, Bessemer, Data Collective

Rocketcrafters State of Florida, DARPA Spinlaunch Adrian Aoun, Asher Delug Stratolaunch Paul Allen

VALT Enterprises Office of Naval Research, Mainte Space Grant

Vector Space Systems Seed Angels, NASA, DARPA, Space Angels, Sequoia Capital Virgin Orbit Virgin Group; Aabar Investments; Saudi Arabia

Zero2infinity Pre sales, Investors, Caixa Capital Table 1. 3 Sources of funding [1]

From the previous surveys and estimates, it is probable that the market will not be able to support most of these new entrants, but it is promising that both the founders and the capital markets project that there will be room for several more participants. Some of the new entrants as discussed has already commenced operations, and a number of other players are likely to have their first flight very soon.

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CHAPTER-2 LITERATURE REVIEW

2.1 Scope – Thesis Research

With the objective of providing a cost-effective launch method, a plethora of techniques is available in the public domain. The next step of the thesis is to analyze multiple methods for their reliability, cost, convenience and efficiency. There are ongoing projects focusing on different types of launch methods already. The table shows different launch types and locations which are under development and those have already successfully been tested. The launch location can be either land, water or Air according to the technique they use.

Vehicle Name Launch Type Launch Location

Arion 2 Land South Europe

Bagaveev Land, Sea

Bloostar Balloon Int'l Water

Boreas-Hermes Balloon

Cab-3A Air KSC, Int'l Water

Chang Zheng 11 Land, Sea China

Chariot Air

Cloud IX Air

Electron Land Birdling's Flat, New Zealand

Eris Land Queensland, Australia

Firefly Alpha Land VAFB, Cape Canaveral, Spaceport

Camden, Wallops

Helios Land

Intrepid-1 Land Kennedy Space Center

Kaituozhe-2 Land China

Kuaizhou-1A Land China

Landspace-1 Land Wenchang, Hainan, China

Launcherone Air Int'l Water

Minotaur I Land VAFB, KLC, WFF, CCAFS

NEPTUNE N1 Land, Sea Moody Space Centre, Australia. Int'l Water;

Orbital 500R Air Malta

Pegasus (Strato) Air Mojave, CA

Pegasus XL Air Int’l Water – Multiple locations

demonstrated

Prometheus-1 Land

PSLV Light Land

RS1 Land Camden, GA; Kodiak, AK

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Skyrora XL Land Scotland

Space Rider Land Kouru

Star-Lord Sea KSC, 20 km offshore

Tronador II Land Puerto Belgrano Naval Base

VALT Land, Sea, Air

Vector-R Land, Sea Kodiak, Cape Canaveral, Wallops

VLM-1 Land Alcatara, Brazil

Table 2. 1 Vehicle launch characteristics [1]

Even though there are studies going on in multiple unconventional methods globally such as, Transport system for space elevator and tether system, Cable space accelerators. Circle launcher and space keeper. Optimal inflatable space towers, Kinetic space towers, Gas tube hypersonic launchers. Earth-Moon cable transport system. Earth-Mars cable transport system. Kinetic anti-gravitator (Repulsator). Centrifugal space launcher. Asteroids as propulsion systems of spaceships. Multi-reflex propulsion systems for space and air vehicles and energy transfer for long distance. Electrostatic Solar wind propulsion. Electrostatic utilization of asteroids for space flight. Electrostatic levitation on Earth and artificial gravity for spaceships and asteroids. Guided solar sail and energy generator. Radioisotope space sail and electro-generator. Electrostatic solar sail and a lot of others, the scope of this thesis is confined to Air launch methods which is proven to be a practically feasible method by past demonstrations. There are different kinds of air launch methods available today. Some of them under research and some tested successfully.

2.2 Lighter than Air techniques

Air balloons, airships gliders etc. Has been in the limelight for being a potential launch method for years, the idea of conducting launches via balloon, known as a 'rockoon,' has been around for a while. Thus far, only sub-orbital launches have been conducted in this manner.

Earlier, balloon had been used as a carrier aircraft [4]_{. The first rocket launch from a balloon} (called Rockoon) was in 1952 from the icebreaker East wind off the coast of Greenland by a research group headed by James Van Allen. It carried a 9 kg payload to an altitude of 95–110 km. The Rockoon (balloon-launched rocket) consisted of a small high-performance sounding rocket launched from a balloon above most of the atmosphere. The Office of Naval Research and the University of Iowa launched 142 Rockoons of four different types in 1953–55 and during the International Geophysical Year 1957–58, from ships in the sea between Boston and Thule, Greenland.

There are currently several teams that have proposed a balloon launch for their X-Prize vehicles. There are several disadvantages to balloons. Balloons must be very large in order to carry anything but the smallest lvs. Launch can occur only on calm days. Recent experience with large balloons during round the world flight attempts has shown that the balloons are always damaged on landing and that they cannot be reused.

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2.3 Internal Air launch from cargo aircraft

Similar to the launch of war missiles, the launch vehicle is stored inside the aircraft as cargo and launched at a certain altitude. A vehicle (up to 39607.6 kg already demonstrated) from an unmodified aircraft and no external indications that the cargo aircraft is carrying a launch vehicle. Propellant boil-off concerns are minimized since the launch vehicle is not subject to either radiation heating from the sun or convective heating from the air stream. The benign environment inside the carrier aircraft allows maintenance and safety problems to be detected prior to launch. An 84,289 lb LGM-30A Minuteman I missile and launch sled was successfully launched on 24 October 1974 from a C-5A Galaxy [11]_.

Figure 2. 1 Lighter than air technique conceptual design

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2.4 Captive on Top

Even though no actual demonstration of this method is not done so far, the space shuttle has used this approach. The concept is logically easy to comprehend. Even though it can carry very large launch vehicle the high cost for modification is a barrier. Active launch vehicle controls are necessary during release to prevent hitting from the aircraft and wings should be large enough to support the launch vehicle at separation from the aircraft, nevertheless, external carriage of the rocket destroys the fuselage lift and causes a large amount of drag that in turn limits launch altitude [5]_.

2.5 Towed launch

Aero-tow is a technique where one aircraft pulls another trailing aircraft horizontally by means of a tether or a stiff connection. This aero-tow method can be likened to a truck pulling a trailer, with a distinct advantage in the amount of mass that can be carried. Easy separation is one of the main benefits of this technique. The need for wings and wheels on the launch vehicle that is sized for takeoff with a full propellant load and the need for a multi-stage launch vehicle with either a pilot or sophisticated flight control system in the 1st stage to maintain position behind the towing aircraft etc are the major disadvantages of towing mechanism. Safety concerns include broken towlines and a

towing aircraft takeoff abort. One of the first occurrences of towing a rocket-powered aircraft was during the summer months of 1942 at Peenemünde, Germany when twin-engine-powered Bf-110C fighters were used to tow prototypes of the Me-163 rocket fighter for flight tests, typically to altitudes of 16,000 ft. Recently NASA successfully flight-tested a prototype, twin-fuselage towed glider that could lead to rockets being launched from

pilotless aircraft at high altitudes—a technology application that could significantly reduce cost and improve the efficiency of sending small satellites into space. The one-third-scale twin fuselage towed glider’s first flight took place Oct. 21, 2014, from NASA’s Armstrong Flight Research Center in California [5]_.

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2.6 Scramjet/Ramjet technology

Nearly 70% of the propellant (fuel-oxidiser combination) carried by today’s launch vehicles consists of oxidiser. Therefore, the next generation launch vehicles must use a propulsion system which can utilise the atmospheric oxygen during their flight through the atmosphere which will considerably reduce the total propellant required to place a satellite in orbit. Considering the strategic nature of air-breathing technology which has the potential to bring a

significant shift in the launch vehicle design, worldwide efforts are on to develop the technology for air breathing engines. Ramjet, Scramjet and Dual Mode Ramjet (DMRJ) are the three concepts of air-breathing engines which are being developed by various space agencies. A ramjet is a form of air-breathing jet engine that uses the vehicle’s forward motion to compress incoming air for combustion without a rotating compressor. Fuel is injected in the combustion chamber where it mixes with the hot compressed air and ignites. A ramjet-powered vehicle requires an assisted take-off like a rocket assist to accelerate it to a speed where it begins to produce thrust.

Ramjets work most efficiently at supersonic speeds around Mach 3 (three times the speed of sound) and can operate up to speeds of Mach 6. However, the ramjet efficiency starts to drop when the vehicle reaches hypersonic speeds.

A scramjet engine is an improvement over the ramjet engine as it efficiently operates at hypersonic speeds and allows supersonic combustion. Thus, it is known as Supersonic Combustion Ramjet or Scramjet [39]_.

2.7 Air-Launch

The fundamental and the most important characteristic considered while formatting most of the space projects is the deltav required. The launch of a rocket from an aircraft or other reusable platform provides a reduction in the required deltav to achieve orbital altitude and velocity.

There are also more benefits other than the initial altitude and velocity of the launch vehicle. Figure 2. 4 Scramjet engine

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Since the conventional ground-launched vehicle initially passes through the densest layers of the atmosphere, there is a significant amount of drag loss involved resulting in the consumption of a major share of the vehicle propellant before it reaches the launch altitude for an air-launched vehicle. The effect of drag on the vehicle diminishes as the atmosphere thins during the ascent. At about 10 km altitude, the density of the atmosphere is only 25% of the density at sea level. As the density of the atmosphere reduces significantly as the altitude increases. At about 10 km altitude, the density of the atmosphere is only 25% of the density at sea level. A diminishing gravity loss is achieved for air launch as the time that an air-launched vehicle needs for the ascent will be shorter than for a ground-launched vehicle. A reduction in steering losses is also possible due to the shorter flight time of an air-launched vehicle [8]_.

From the literature review, multiple results can be stated supporting the argument of choosing Air launch as the most effective method.

• Air launch provides a reduction in DV to orbit of ~300-950 m·s-1 for subsonic launch conditions.

• The optimal release flight path angle is around 30° but has for subsonic release velocities a range of almost ±15°.

• There is no need for a wing for close to optimum release flight path angles, however, for small release flight path angles a wing can reduce the required DV with ~100-200 m·s-1 compared to a wingless launch vehicle

• A wing might also be required for the pitch up since aerodynamic control surfaces are more effective than thrust vectoring control (TVC)

• Air launch more beneficial for small launch vehicles since drag loss is more significant for such vehicles.

• A more efficient nozzle design can be utilized for an air-launched vehicle due to the lower ambient pressure at launch altitude.

• Can provide a more convenient design altitude (ambient pressure equals the exit pressure of the nozzle and ideal expansion will occur) since it operates over a smaller range of pressures.

• As the density is reduced, the loads on the rocket are minimal and the structural design can be simplified as the dynamic pressure depends on the atmospheric density and velocity of the vehicle.

• Reduction in acoustic loads compared with the ground launch. Acoustic reflections from the ground can damage the launch vehicle and often requires additional structural reinforcements for the launch vehicle.

• Since an air-launched vehicle does not require a fixed launch site, a wide range of orbital Inclinations can be utilized due to the mobility of the launch platform

• The flexibility of launch latitude and azimuth reduces the expense, in terms of DeltaV, dog-leg maneuvers etc.

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• Air launch is not restricted to the extremely demanding weather conditions that are imposed to a ground launch as the high altitude can be adjusted as to avoid unfavorable climatic situations.

One main challenge to be addressed practically is the certification, the ignition of a rocket engine in the proximity of a manned aircraft is a dangerous scenario and might require special clearance.

2.8 Parameters affecting velocity calculation.

The study Air-Launching Earth to Orbit: Effects of Launch Conditions and Vehicle Aerodynamics in the Article in Journal of Spacecraft and Rockets · May 2005 11 studied different aspects such as carrier aircraft specifications, Velocity calculations, Launch conditions, Vehicle Aerodynamics, Altitude, Flight-Path Angle, characteristics of wings etc[43]_.

A launch vehicle must provide a change of velocity, V or delta-V, in order to deliver a payload to low Earth orbit (LEO). The delta V from the launch vehicle’s propulsion system depends on the time histories of the vacuum thrust and the mass of the LV. Videal from the LV’s propulsion system must equal the Videal required for an earth to orbit trajectory as follows:

∆Videal = Vorbit + ∆Vdrag + ∆Vgravity + ∆Vsteering +∆Vatmospheric pressure − Vearth rotation −

Vcarrier-aircraft

The orbit Vorbit and Earth rotation Vearth-rotation velocities are driven by the satellite requirements, while the velocity losses Vdrag, Vgravity, Vsteering, Vatmospheric-pressure are driven by the trajectory that connects the launch point to the orbit. The carrier aircraft’s flight velocity at separation Vcarrier-aircraft is a function of the carrier Vcarrier-aircraft capabilities and the release conditions.

Some assumptions are made while formulating the equations such as,

1.The change in velocity caused by Coriolis acceleration is ignored since the magnitude is very small, which is less than 10 m/s.

2.Any wind occurring during the flight is assumed to be zero at all altitudes.

3.The first component Vorbit is the payload orbital velocity, and it depends on the altitude of the perigee and apogee of the orbit.

4.Drag losses Vdrag are caused by friction between the launch vehicle and the atmosphere and are on the order of 40–160 m/s for medium-sized launch vehicles such as Delta or Atlas rockets for a ground-launched Earth-to-orbit trajectory.

5.Gravity losses Vgravity arise because part of the rocket engine’s energy is wasted holding the vehicle against the pull of Earth’s gravity. They are highly dependent on the thrust-to-weight (T/W) ratio and are on the order of 1150–1600 m/s for ground-launch vehicles.

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6.Steering losses Vsteering are caused by the need to steer the launch vehicle. Atmospheric pressure losses Vatmospheric-pressure are the difference in performance of a rocket motor in a vacuum as compared to its operation in the atmosphere [43]_.

The best performance a rocket motor can provide is in a vacuum. Rocket thrust is calculated from the momentum change of the fuel and the difference in pressure at the nozzle exit. Any ambient pressure reduces the thrust of the rocket motor in the amount of atmospheric losses. Air launching always reduces atmospheric pressure losses because ambient pressure is lower at altitude as compared to sea level. The Earth rotation velocity increment Vearth-rotation depends on launch latitude and launch direction. Finally, the carrier aircraft’s flight speed Vcarrier-aircraft directly reduces delta V required from the launch-vehicle’s propulsion system.

The advantages from the carrier aircraft’s flight speed is mostly dependent on both the T/W ratio at release conditions and the flight-path angle that is the angle between the launch-vehicle velocity vector and the local horizontal [43]_.

2.9 Launch condition and launch vehicle characteristics

They used an existing launch vehicle, the Minotaur LV, to check the numerical results with actual flight data. With the exception of its aerodynamic data, its characteristics are published. They simulated air launching the Minotaur over a range of altitudes, launch speeds, and launch flight-path angles. The altitudes chosen are sea level through 30,490 m at every 7620 m. The velocities are Mach 0, 1, 2, and 3, taken at sea level. The launch velocities were pushed just beyond current state-of-the-art speed of Mach 3 in order to capture some of the current proposed methods of launch. The launch flight-path angles are from horizontal to vertical, incremented at every 30 deg. Each launch condition was simulated with and without a wing. More than 160 simulations were conducted in which launch altitude, speed, and flight-path angle were varied, and the effect of adding a wing was also modeled. The Minotaur is a small 36,200-kg four-stage solid rocket that is currently ground launched. It consists of the first two stages of the Minuteman II and the upper two stages of the Pegasus XL. The Minotaur has its initial T/W ratio relatively high at 2.3 to 1, which is desired for an air-launched launch vehicle to reduce losses.

Three methods were used in determining the accuracy and sensitivity of the aerodynamic data. The payload performance, payload sensitivity, and velocity losses are all evaluated and compared to published values. The vehicle payload performance is evaluated by comparing two different launch simulations to the published values. The two launches were a 400 n mile sun-synchronous orbit from California and a 100 n mile 28.5-deg circular orbit from Florida. Both launches give a performance measure for comparison with the simulations. The simulation payload performance was very close to the published values with the percentage differences are at +5.6 and −0.67% respectively. The payload sensitivity was evaluated via varying the drag coefficients and running the POST simulation to find the resulting payload. Estimated data were obtained from Datcom program.

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Payload only changed by ± 8% when drag coefficients were varied from 25 to 200%.

Delta-velocity gain Vgain was obtained by taking the difference between the delta V calculated by POST for a ground launch from Cape Canaveral Vground-launch to a 185-km 28.5-deg inclination circular orbit and the delta V for an air launch Vair-launch from the same location to the same orbit. Trajectory optimization program calculated that the reference ground launch orbit had a delta-V budget of 8918 m/s. The Minatour launch can place 646 kg in the orbit from Florida with total losses of 1534 m/s.

Delta-V gain from air launching is an improvement as compared to the ground launch as Vgain = Vground launch − Vair launch

The Vgain results are organized into two sets of plots as no wing and wing for pull up. Positive values on the vertical axis indicate an improvement in Vgain, and negative, a reduction in performance. The horizontal axis is the Launch vehicle’s flight-path angle at engine ignition. The Launch vehicle flight-path angle at ignition will normally be less than the carrier aircraft flight-path angle because the Launch vehicle engines are typically started sometime after release for the safety of the carrier aircraft. Also, the Launch vehicle’s altitude and airspeed will be typically less than the carrier aircraft’s. Both Figures are based on Launch vehicle flight-path angle, airspeed, and altitude at engine ignition and not on the carrier aircraft’s parameters at Launch vehicle release. Also, the Launch vehicle’s body axis is aligned with the Launch vehicle’s velocity vector at engine start.

2.9.1 Altitude effect

Multiple trends are evident on these plots regarding the altitude. All of the plots show that an increase in altitude increases Vgain. The primary reason for the reduction in atmospheric density. The greater thrust that a rocket engine provides at altitude allows the vehicle to accelerate faster, which in turn decreases the time that the vehicle is steering and fighting gravity. Hence all four major losses decreased with an air launch

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2.9.2 Effect of flight path angle.

The optimum flight-path angle for best Vgain spans a range of about ±15 deg because the curves in Figures are relatively flat near their peaks. Adding a wing reduces the optimum flight-path angle for maximum Vgain at low angles and low carrier aircraft speeds. The optimal launch angle is different for each launch condition. The optimal flight-path angle can best be described as the trajectory with the least maneuvering. Pull ups and other sudden maneuvers increase steering and drag losses. Air launches at a low initial flight-path angle and at a low altitude

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cause the Launch vehicle to pull up in order to fly the optimum trajectory to orbit. This pull-up causes increased losses. At high altitudes, too steep of an initial flight-path angle results in the Launch vehicle pitching down in order to fly the optimum trajectory. Pitching down increases the losses. A minimum flight path of 30 deg above the horizontal at ignition can be used for a first approximation to maximize Vgain. A release attitude of less than 30 deg might require a wing.

2.9.3 Effect of the addition of wings on Launch Vehicle.

The advantage of a wing is the aerodynamic lift it can provide. The wing provides the ability to turn more efficiently, but it only provides benefits at launch when the launch angle is near the horizontal. In an expendable launch vehicle, to maximize the effectiveness of the wing it should be used only for the initial pull-up maneuver and then immediately jettisoned to reduce weight. The wing reduces the angle of attack required to complete a pull-up maneuver, which in turn reduces drag. Without the wing, the turn is completed by using engine thrust while the rocket’s body briefly generates large amounts of drag. If the carrier aircraft releases an Launch vehicle at a flight-path angle that is close to the optimum.

In such case the wing provides no Vgain. A wing can help if the carrier aircraft cannot pull up to the optimum flight-path angle.

2.9.4 Effect of the carrier aircraft velocity.

The carrier aircraft’s flight speed Vcarrier-aircraft improves Vgain. The Vgain from air launching can exceed the contribution made by the carrier aircraft’s flight speed or the carrier aircraft’s flight speed might not even provide a one to one benefit toward Vgain. The magnitude of the benefit depends on the Launch vehicle’s flight-path angle at engine start.

Based on the estimation, the most beneficial carrier aircraft launch parameters are in the following order: launch velocity, launch flight-path angle, and launch altitude. In addition, there is an optimum launch flight-path angle that maximizes the velocity benefit from air launching.

2.10 Candidate launcher features:

The launch vehicle definition is very basic and considerable assumptions are made taking different existing launch vehicles in to account. From the literature review, it is understood that the liquid rocket motors have considerable advantages for the air launch mechanism, even though the solid rocket motors are theoretically dominant. A detailed study of different similar launch vehicles which are under development and those which are already proven such as Pegasus is considered and basic design of the launch vehicle is prepared.

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Some common commercial engines are chosen according to the thrust and weight requirements and the selected ones are described later in this report. From the brief literature review, a generic dimension and an accumulated weight is calculated and the pylon design is analyzed so as to fit the needs [26]_.

2.11 Candidate Aircraft features:

Being a project aimed at acquiring commercial success and profit, A number of practical requisites have to be met. A commercial aircraft available to any normal entrepreneur is one of the prime conditions while choosing the aircraft. The large payload capacity, modification cost, maintenance cost etc. Are also to be considered during selection. A detailed analysis of the financial aspect of the project is also done using available research methods later in the project stage. Currently, existing studies done by the DARPA mission is analyzed to provide us with a detailed solution to start the design. Multiple existing carriers used previously for similar or significant missions are subjected to detailed analysis in the study and some conclusions are drawn [38]_.

The Antanov An-225 which is a Ukranian built airlift cargo aircraft used to ferry the Soviet Buran orbiter, Boeing 747-100 used for carrying the space shuttle orbiter, Dual fuselage variant of the c-5 galaxy strategic airlift carrier aircraft, the white knight xx, A-380, 747-400 etc are the main candidates taken into consideration for the study. The main parts of the results of the DARPA study are tabulated as following and the final decision are explained in the system concept definition.

Carrier aircraft Cases analyzed Cases with positive payload

Launch vehicle gross weight available (kg) Maximum LEO payload (kg) White Knight X 195 180 79832.2 5071.2 747-100 (SCA) 195 185 108862.2 7003.5 A380 195 184 119997.9 7751.9 747-400F 195 185 139706.4 9071.8 An-225 195 185 200000.2 13780 White Knight XX 195 185 340194.3 22652.4 Dual-fuselage C-5 195 185 350000.0 23718.3 Total 1,365 1,289

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Stage 1 Stag e 2 Stag e 3 Paylo ad (kg) Recurri ng cost/flig ht ($) Lifecyc le Cost/k g ($) Total Faciliti es ($) Total DDT &E ($) Total producti on ($) Total Operati ons ($) Total lifecyc le ($) Probabili ty Of LOM (%) White knight X Drop tank/LH2 LH 2 5071. 2 113 11540 91 1470 12430 1460 15460 1.9 Solid Soli d 1832. 5 21 7360 77 660 2270 550 3560 2.1 Solid Soli d Soli d 2417. 6 27 6490 89 760 2620 680 4150 2.4 747-100 SCA-911 Drop tank/LH2 LH 2 7008 136 9860 110 1550 14790 1800 18250 2.6 Solid Soli d 2517. 4 24 5880 94 640 2490 690 3910 2.7 Solid Soli d Soli d 3311. 2 31 5240 110 760 2870 840 4590 3.1 747-400 F Drop tank/LH2 LH 2 9071. 8 157 8860 129 1860 17140 2090 21230 2.8 Solid Soli d 3243. 2 26 5320 112 860 2800 770 4550 2.9 Solid Soli d Soli d 4259. 2 34 4730 132 1000 3230 950 5310 3.3 A380-800F Drop tank/LH2 LH 2 7751. 8 144 9630 117 1630 15980 1920 19650 3 Solid Soli d 2775. 9 25 6130 101 680 2930 730 4440 3.1 Solid Soli d Soli d 3655. 9 32 5400 118 810 3330 890 5150 3.5

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2.12 Thesis work flow chart

Market study

Window for a new launch vehicle Demand and financial feasibility Structural design Mission design Analysis of innovative methods

Modal Analysis Buckling/Fracture Static structural/ mechanics analysis

Modify Design

Final Design Result and findings

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CHAPTER-3 SYSTEM CONCEPT DEFINITION

Considering the extensive literature review and the global market study carried out in the first phase of the thesis, the basic assumptions and ground rules are to be listed. A suitable launcher-aircraft combination has to be selected from the available data. Assumptions are made also by taking practicality into account. Theoretical and classical methods are used to define the various technical specifications of the launcher. The data from similar projects which are under development is also used to compare with the results obtained from the calculations and verify them for their feasibility.

3.1 Carrier aircraft specifications.

From the analysis, 747-400 was selected as the candidate aircraft as the other aircrafts were eliminated. White Knight X was eliminated due to its low payload capability compared to existing commercial aircraft. An-225 had multiple disadvantages since only one unit is available currently, and the risks of purchasing and maintaining such a unique specimen were very high considering this is a commercial project. A380 and 747-400 have similar specifications but A380 being more expensive, it was also eliminated.

The calculated optimized gross weight of the launch vehicle by using existing rocket motor is below the maximum capability hence reducing the structural modification required resulting in lower development cost. The availability of an extra adapter for the pylon which is originally designed for additional engines can be used for the wing carrying structure which makes 747-400 a better selection.

The aircraft is of wide-fuselage, low-wing design with four podded underwing turbofan engines. Optional engine fits include Pratt & Whitney PW4062, General Electric CF6-80C2B5F and Rolls-Royce RB211-524H2-T turbo-fan engines, developing between 252kn and 276kn. There are four main fuel tanks in the wings, a centre wing tank, a tailplane tank and reserve fuel tanks in the outer wing sections. The maximum fuel capacity is 216,846l. The auxiliary power unit is installed in the tail section. Even though there exists no marginal difference which has a significant effect on our specific mission, the general electric engine is selected for the system design considering the commercial advantages [6]_.

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Aircraft type Boeing 747-400, Boeing 747-400 Freighter Manufacturer Boeing Commercial Airplane Company Type of operations Medium to long range widebody airliner Dimensions Length 70.66 m

Wingspan 64.44m Height 19.33m Wing reference area 510.9667m2 Wing weight 43,090 kg Wing sweep 37.50 Engine types Four General Electric CF6-80C2B1F

Four Pratt & Whitney PW4056 Four Pratt &Whitney PW4062 Four Rolls Royce RB211-524H Take-off thrust options TO (Full thrust)

TO1 (5% thrust reduction) TO2 (15% thrust reduction) Take-off flaps settings FLAPS 10

FLAPS 20

Take-off runway conditions DRY 2mm (0.08in) Water 6mm (0.25in) Water 13mm (0.50in) Water Landing runway conditions DRY

WET

CONTAMINATED Table 3. 1 747-400 specifications [16]

Aircraft type Boeing 747-400 Boeing 747-400F

Cargo capacity (Max) 60400 kg 121729 kg

Usable Fuel Capacity (Max) 173474 kg 164111 kg

Dry operating Weight 178756 kg 165393 kg

Max Zero fuel weight 246074 kg 288313 kg

Max take-off weight 396894 kg

Max Landing weight 285764 kg 302093 kg

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General Electric engines Basic fuel capacity Maximum fuel capacity (1 body tank) (2 body tanks)

Passengers (FC/BC/EC) 416 (23/78/315)

Cargo pallets 4/14 3/14

Engines CF6-80C2B5F CF6-80C2B5F

SL standard-day takeoff thrust/flat-rated temperature

(BET) lb/°F 62,100/86 62,100/86

Maximum taxi weight kg (lb) 414,130 (913,000) 414,130 (913,000)

Maximum takeoff weight kg (lb) 412,760 (910,000) 412,760 (910,000)

Maximum landing weight kg (lb) 263,530 (581,000) 295,740 (652,000)

Maximum zero fuel weight kg (lb) 245,840 (542,000) 251,740 (555,000)

Operating empty weight3 _{kg (lb)} _185,610 _{(409,200) 186,330} _(410,800)

Fuel capacity L (U.S. gal) 228,160 (60,275) 240,310 (63,485)

Design range (MTOW, full passenger payload) nmi (km) 7,4954 _{(13,880) 7,565} _(14,010)

Cruise Mach 0.855 0.855

Takeoff field length (SL, 86°F, MTOW) m (ft) 3,320 (10,900) 3,320 (10,900)

Initial cruise altitude (MTOW, ISA + 10°C) ft 31,900 32,800 Landing field length

(MLW) m (ft) 1,935 (6,350) 2,175 (7,150)

Approach speed (MLW) kias 147 157

Fuel burn/seat

6,000 nmi kg (lb) 304.5 (671.4) 305.4 (673.4)

Table 3. 3 747-400 GE engine characteristics

3.2 Flight envelope

The statistical and technical data of the candidate aircraft is studied, and the basic aerodynamic and structural coefficients are noted. The aircraft envelope, dimensions, normal maneuvers performed, normal takeoff procedures, velocity etc. Are taken into account while calculating the load acting on the pylon while on the operation for doing different analysis after the basic design [29]_.

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Flight Phase Description

1) Engine Start and Taxi Engines start and taxi to runway

2) Take off Accelerate to 155 knots (178 mph) to takeoff

to minimum ground clearance altitude (35 feet)

3) Initial Climb Climb to 1500 feet (speed approx. 191 mph)

4) Climb to Cruising Altitude Climb to 35,000 feet (290 knots or 288 mph) with angle of attach reduced to 10 degrees at 16,500 feet

5) Cruise Flight to Release Point (prelaunch checkout)

Accelerate to and maintain 540 knots (621.4 mph) and perform launch vehicle prelaunch checkout

6) Launch Vehicle Release Release launch vehicle on clearance from range and launch vehicle control

7) Cruise Flight to Landing Field Return to airfield (same as departure field in this scenario)

8) Descent Final Approach Descend to 1500 feet and reduce speed to 250 knots (288 mph)

9) Final Approach and Landing Decelerate and land at 135 knots (155 mph) slowing to taxi speed

10) Taxi and Engines Shut Down Taxi to parking ramp and shut down engines

Table 3. 4 Mission phases for Boeing 747-400 carrying air-launched rocket

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3.3 Launcher specifications

A basic study of the different available launchers is carried out, Considering the possibility of using a new engine configuration and technology. The estimated orbit conditions, Maximum payload, Cost, availability etc. Are the major constraints in choosing a launcher. Different technologies are under development as this thesis is under process. The basic assumptions and mission objectives are used to select from a pool of multiple currently available and under development in the market. Few of the candidate launcher configurations are Electron, Launcher-One, Pegasus, thunderbolt, Electron is a small liquid-fueled orbital launch vehicle, which is being developed by Rocket Lab, a US/New Zealand company. Both stages are RP-1 and LOX fueled. The electron has a carbon composite structure of 1.2 m diameter and 20m length.

The power plant of Electron is Rocket Lab's Rutherford engine using LOX/RP-1 as fuels [35]_.

Rutherford adopts a new electric propulsion cycle, using electric motors to drive its pumps, and is the first oxygen/hydrocarbon engine to use 3D printing for all primary components. Electron’s first stage is powered by multiple Rutherford engines with a total peak thrust of 146.6kn. Electron’s second stage is powered by a Rutherford Vacuum Engine which is tailored to provide improved performance in vacuum conditions. An additional kick stage powered by Rocket Lab's Curie engine running on an unspecified “green” monopropellant is available for increased flexibility of the missions.

Launcher-One [20]_{is a two-stage air-launched vehicle using two Virgin-designed and built}

Newton RP-1/LOX liquid rocket engines. The rocket has a diameter of 1.6m for the first stage

and 1.3m (4ft 3in) for the second stage and payload fairing. The first stage uses one Newton-Three engine, while the upper stage uses one Newton-Four engine. Newton-Newton-Three will generate 330 kilonewtons of thrust while Newton-Four will deliver 22kn (4,900lbf) to the second stage. It is designed to launch "smallsat" payloads of 300 kilograms into Sun-synchronous orbit, Pegasus consists of three solid-fuel stages with an optional HAPS (Hydrazine Auxiliary Propulsion Stage) monopropellant fourth stage [26]_.

For the thesis work, A combination of data from all the above projects is inherited to make a design suitable for the mission objective. The dimension and the specifications used for the thesis work is depicted in the following section. A 3D model of the launcher with the appropriate dimension is created in order to obtain the aerodynamic and classical loads acting on the pylon design.

3.4 Launcher measurements Stage 1

Stage one consists of a Newton 3.1 turbopump-fed LOX/RP-1 booster engine that delivers

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Stage 2

Stage two consists of a Newton 4 LOX/RP-1 booster engine that delivers 22,241N of thrust and Isp value of 328, 1.5m outer diameter composite structure.

Control surfaces

Multiple types of control surfaces were subjected to study including winged launchers such as Pegasus which eases the control mechanism and allows for an easier separation technique. Since there are multiple disadvantages for a winged structure such as additional weight, Complex designing process, Space constraints to fix on the wing which led to a choice of simpler control surfaces using 4 fins as conventional launchers. Advanced electronic systems might be required while launching without wings, which are neglected and assumed to be solved while integration and deployment of the booster.

Payload module

The payload module consists of a conical adapter and is assumed to have an aerodynamically optimized nose cone design. The classical design methods in order to occupy the combination of Single Primary Payload, Small Primary and Multiple cubesat Payloads, Multiple, Similar Primary Payloads (Axial Release), Multiple, Similar Primary Payloads (Lateral Release) etc[23]_.

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3.5 Mission design

The Mission design consists of three major elements: the launch vehicle, carrier aircraft, and the ground segment. The carrier aircraft that carries the launch vehicle under the port (left) side wing between the fuselage and inboard engine. The carrier aircraft provides electrical power, purge gasses, health monitoring, and control of the launch vehicle by a launch engineer onboard the aircraft. The ground segment consists of equipment to load propellants on the launch vehicle, ground stations to gather and distribute telemetry, and a mission control centre to monitor the launch operation.

The fuel management system is a significant department for this project as it stabilizes the aircraft when there is an extra weight of the launcher on only one wing. The propellant has to be stored on the right wing which is of an equal amount of the launcher and the distribution has to be taken care of dynamically. The fuel stored on the right wing will be flown to the left side just after the launcher has been released from the aircraft, this method is commonly used for fighter jets and bombers while dealing with missiles, even though this method has never been used for a large scale it is not considered to be a difficult process as the technology readiness level is considered. The pneumatic and electronics systems engaged in this procedure is not dealt with in this thesis as it beyond the scope.

The launcher as defined is a two-stage vehicle, with a 1.6m diameter first stage and a 1.3m diameter second stage and fairing, constructed primarily of composites to minimize weight and maximize propellant mass fraction. The first stage is powered by a Newton 3.1 turbopump-fed LOX/RP- 1 booster engine that delivers 266,893N of thrust the second stage is powered by a Newton 4 engine that delivers 22,241N of thrust. The vehicle can deliver over 500 kg to a low inclination, low altitude orbit. The maximum take-off weight of the Launcher is assumed to be 25,000kg

The payload assembly is assumed to enable the payloads to be mated with a payload adapter and encapsulated in the payload fairing independently of the first and second stage launch vehicle integration. This allows integration to happen in parallel and allows for pre-encapsulation of payloads if desired by the customer.

The Launcher aircraft with the launcher onboard can take-off from multiple launch locations around the world.

The aircraft will travel for 30 minutes to four hours to reach the required altitude and the deployment zone of the payload. It will release the launcher from its pylon at an altitude of 35,000ft, with an angle of 27° skywards.

After the release, Launcher makes a free-fall descent for four seconds before the first stage engine Newton-Three ignites to life. The engine is assumed to have a run time of three minutes and will propel the rocket at a speed of 12,874km/h.

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The second stage engine will be ignited when the rocket reaches an altitude ranging between 498km and 1,199km. It burns multiple times for six minutes and will propel the rocket at a maximum speed of 28,163km/h.

The fairings open up when the payload reaches its destination and the satellite is placed into its target orbit. It can launch satellites between a 0oand 120orange of orbital inclinations while placing them into low, medium and high Earth orbits, based on the requirements.

Performance Criteria Mission design

Payload capacity 300 kg to 500 km SSO

500 kg to 230km SSO

Release altitude 35000 feet

Release Velocity 0.8 Mach (987.84kmph)

Assumed Precision Apogee/Perigee: +/- 15 km

Inclination: +/- 0.15o RAAN: +/- 0.2o

Launch site According to customer requirements

Table 3. 5 Assumed Performance criteria

3.6 Launcher release mechanisms

Launcher release mechanism is a complex and important part of the mission design, there are multiple methods employed to achieve this since the mechanism is similar to that of launching a missile for weaponry. Some of the most used techniques are Gravity launch systems, Impulse launch systems, Reaction launch systems etc. [22]_.

In Impulse launch mechanism a force is applied to the launch vehicle either to project it along the entire path of its trajectory or, to a lesser degree, to clear the launch vehicle from the delivery vehicle. Two basic launcher types employ this method: cannon launchers and ejector launchers Cannon launchers are more common and may be further subdivided into guns, howitzers, and mortars. The basic differences among these subdivisions are trajectory, initial velocity, and size of the launcher, with guns having the highest muzzle velocities and flattest trajectories. Ejector Launchers are Impulse launchers for weapon ejection which are employed for both free-fall and self-propelled weapons. Their main purpose is to ensure that the weapon safely clears the delivery vehicle. Ejection is usually accomplished by the expansion of high-pressure gases from a compressed air supply or from ignition of a propellant charge. Because it is used for ejection purposes only, the impulse is small, and the launcher can be built to withstand the shock of launching without the need for excessive structural strength or special devices. Thus, launchers of this type are fairly light and simple in design.

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This third general type of launcher is Reaction launchers, the one in which the force separating the weapon from the launcher is contained within the weapon. These weapons are normally rockets or missiles. The propulsion system of the missile itself may be used to provide the necessary force. Thus, most self-propelled weapons, if not launched by ejection, are put into flight by reaction launchers.

Reaction launchers provide static support for the weapon [36]_{and initial flight orientation. They} are characteristically small and light since they are not required to sustain large moments of force upon weapon launch. Reaction-propelled weapons often depend upon wings or fins to provide lift and must use rocket thrust to overcome gravity temporarily and to propel the weapon to desired flight speed. If, during the development of thrust, a weapon is free to move along the launcher, it might not have sufficient thrust or lift to overcome gravity at the time it leaves the launcher. Thus, the missile could fall to the deck of a launching ship or become completely disoriented before sufficient thrust or lift had been developed to sustain its flight. To prevent this from happening, the weapon is restrained on the launcher until sufficient thrust is generated. The restraining device may be simply a pin that is sheared when the weapon develops the required thrust, or it may be a more complicated, reusable device that releases the weapon when the required thrust is exerted.

The term rail launcher may be applied to launchers making use of rails, tubes, long ramps, and even tall vertical towers. All provide, to a varying degree, constraint to the weapon while it is moving on the launcher, and they thus provide a considerable amount of flight control. For uncontrolled weapons, such as rockets, the rails must be fairly long so that the rocket is constrained for a longer portion of the rocket motor burning time to provide the necessary initial velocity vector control. If the missile is equipped with a guidance system, the rail length can usually be reduced. Long-range weapons, guided or unguided, normally require a longer rail since their initial acceleration is comparatively low relative to short-range weapons. A rocket booster may be employed, however, to provide sufficient acceleration to permit reduction of rail length.

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CHAPTER – 4 LOAD ESTIMATION

4.1 Design and Load estimation

As this thesis covers only the preliminary analysis, a number of assumptions are made, and a very limited amount of inputs are available as the inputs are unlikely to be available for commercial projects since they are confidential and classified. Most of the calculated load, dimensions, material properties, cost etc. Are purely based on theoretical and statistical data. The different materials used, Aerodynamic loads calculated, first approximation of the design, Weight estimation, supports considered, reaction forces calculated with the classical approach, Pylon integration structures suggested etc. Are discussed in detail in this chapter.

4.2 Design

The preliminary pylon design needs to take multiple loads into account such as the weight of the launcher, The drag from the launcher, The drag of the pylon, the reaction load from the support to the wing structure. A basic design of the pylon is created considering different literature and studies conducted by DARPA [19]_{, ALSAT} [22]_{, ISRO} [40]_{, Virgin group} [20]_etc. With approximated dimensions and a classical approach with lateral and longitudinal support structures. The basic design is depicted in the figure

A preliminary design of the supporting structures such as the pylon hooks, front and rear fittings are also generated via CATIA V5 for an understanding of the preliminary dimensions which are under consideration.

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Figure 4 2 Front fitting, 2. Hook, 3. Rear fitting, 4. Sway Brace

4.3 Weight estimation:

As discussed in the system concept definition the weight of the launcher is approximated to be 25000kg, Considering the flight envelope of the carrier aircraft in the event of performing a maneuver the load can be up to 2.5 times the weight and hence in order to do the static structural analysis, the weight of 2.5 times 25000 Kg is applied. One major assumption took for this study is that the center of Gravity is always aligned for all three components, i.e. The wing, pylon and the launch vehicle. As a preliminary analysis, only the static loads are considered, and no dynamic simulations are carried except for the drag estimation. The weight of different support structures such as the pylon hook, sway braces, rear fittings etc. Are neglected since they are

1 2

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comparatively lightweight structures. The takeoff weight may vary depending on the mission objectives such as the orbit, Delta V, payload weight, propellant weight etc., and hence the maximum values are always considered.

Since the pylon integration structures such as the sway brace, rear fittings etc. Are not clearly validated, the launch vehicle weight is assumed to be distributed equally on the bottom surface for the modal as well as static structural and buckling analysis.

4.4 Drag

Drag is a force that opposes the direction of motion of a vehicle through a fluid. In terms of atmospheric effects on a launch vehicle, drag has the most significant negative effect on the performance. There are many different types of drag, and all act in a way to impede the motion a vehicle. The drag types considered for the launch configurations are zero lift drag, drag due to lift, friction drag, pressure drag, and base drag. These drag components on a launch vehicle can be broken up into three classes. The first class is simply labelled drag and for this project consists of zero lift drag coefficient (CD0) and lift-induced drag coefficient (Ci). Zero-lift drag is the total drag of the vehicle when it has the condition of zero lift, and drag due to lift is the coefficient of the vehicle when it is not at a zero-lift condition. The total drag coefficient (Cd) is the summation of these two components. The equation for evaluating drag is very similar to lift, drag (D) is also dependent on air density (ρ), cross-sectional area (A), and velocity (V) relative to the air. The difference between the two equations comes with an estimation of the drag coefficient. The direction of drag calculated in this equation is always opposite of the vehicle’s velocity vector [10]_.

𝐷 = 1 2𝜌𝑉

2_𝐶 𝑑𝐴

Since the pylon deals with only the drag while the launch vehicle is attached to the pylon until its release, forces such as Pressure Drag, Base Drag, Aerodynamic Heating, Aerodynamic Acoustics, Static Stability etc. Are neglected.

In order to calculate the density property, an atmospheric model of the atmosphere has to be defined, Atmospheric modelling is an important method to generate physical and numerical measurements of climate parameters, quantify the spatiotemporal changes of atmospheric phenomena over space and time, and predict their occurrences. We use the GFS model for obtaining the necessary data. The Global Forecast System (GFS) is a weather forecast model produced by the National Centers for Environmental Prediction (NCEP). Dozens of atmospheric and land-soil variables are available through this dataset, from temperatures, winds, and precipitation to soil moisture and atmospheric ozone concentration. The entire globe is covered by the GFS at a base horizontal resolution of 18 miles (28 kilometers) between grid points, which is used by the operational forecasters who predict weather out to 16 days in the