Rocket Engines and Launchers PDF

© ESA Rocket Engines and Launchers Jordi L. Gutiérrez Department of Physics (UPC) Institute of Space Studies of Catalonia Introduction to Space EETAC, Fall 2024–2025 […] as I hur...

© ESA Rocket Engines and Launchers Jordi L. Gutiérrez Department of Physics (UPC) Institute of Space Studies of Catalonia Introduction to Space EETAC, Fall 2024–2025 […] as I hurtled through space, one thought kept crossing my mind – every part of this rocket was supplied by the lowest bidder. John Glenn, US Astronaut Rocket science is tough, and rockets have a way of failing. Sally K. Ride, US Astronaut Most of the time, [for the customers, the launcher] will be a very expensive nuisance, and one of the greatest risks to the success of [their] mission. James S. Wood, in Space Mission Engineering: The New SMAD.  Syllabus 1. Physical basis 2. The rocket engine 3. Step rockets 4. Modern launchers 5. Launch bases 6. The near future  At the end of the day you should know 1. How rocket engines work. 2. Rocket’s figures of merit. 3. Some pros and cons of solid motors an liquid engines. 4. Why do we need stage rockets as launchers.  © Quino PHYSICS FUNDAMENTALS  Required Physics Rocket engines theory is based on Classical Mechanics (1B), Chemistry (1B), Thermodynamics (2A), and Fluid Mechanics (2B). If we consider electromagnetic engines, we must add Electromagnetism to the list. You will learn in earnest on the subject Rockets and Launches (3A). © Physics Stack © SpaceX  Newton’s Principles Recall that these principles are only correct when using an inertial (non-accelerated) reference frame. First principle: an object subjected to a null net force remains in a state of motion with constant velocity. Second principle: a net force applied to a body causes a change in its linear momentum, 𝑑𝑑 𝑝𝑝⃗ 𝐹𝐹⃗ = with 𝑝𝑝⃗ = 𝑚𝑚𝑣𝑣⃗ 𝑑𝑑𝑑𝑑 Third principle: if a body 𝐴𝐴 applies a force upon a body 𝐵𝐵, this latter body applies a force upon 𝐴𝐴 such that 𝐹𝐹⃗𝐴𝐴→𝐵𝐵 = −𝐹𝐹⃗𝐵𝐵→𝐴𝐴  Embedded question 1 After publishing A Method to Reaching Extreme Altitudes1, Robert H. Goddard was publicly criticized (scorned?) in the New York Times: «That professor Goddard, with his 'chair' in Clark College and the countenancing of the Smithsonian Institution does not know the relation of action to reaction, and of the need to have something better than a vacuum against which to react –to say that would be absurd. Of course he only seems to lack the knowledge ladled out daily in high schools.» A retraction was published by the NYT on the 17th of July, 1969, almost 25 years after Goddard’s death. Can you justify why such a retraction was necessary? You have 5 minutes. 1. There is an affordable edition published by Dover Ed.  © NASA GSFC Every vision is a joke until the first man accomplishes it; once realized, it becomes commonplace. Robert H. Goddard  Conservation of linear momentum The second principle states that if no net external force acts upon a system, it must conserve its linear momentum: 𝑝𝑝⃗i = 𝑝𝑝⃗f with 𝑝𝑝⃗ = 𝑚𝑚𝑣𝑣⃗ where 𝑝𝑝⃗i and 𝑝𝑝⃗f are, respectively, the initial and final linear momenta of the system. If the system consists of 𝑁𝑁 particles, then 𝑁𝑁 𝑁𝑁 𝑝𝑝⃗i = 𝑝𝑝⃗i,𝑘𝑘 = 𝑝𝑝⃗f,𝑘𝑘 = 𝑝𝑝⃗f 𝑘𝑘=1 𝑘𝑘=1  Embedded Problem 1 A person of mass 𝑚𝑚 is on a wagon with mass 𝑀𝑀 which carries 𝑁𝑁 bricks, each one with a mass of 𝑚𝑚b. The person takes one brick and throws it away with a perfectly horizontal velocity 𝑣𝑣b. Determine the velocity of the system © Arthur C. Clarke (person + wagon + remaining bricks). What will be the velocity of the system person + wagon after all the bricks have been ejected?  Total Impulse From the second principle, we know that 𝑑𝑑𝑑𝑑 𝐹𝐹 = 𝑑𝑑𝑑𝑑 then 𝑑𝑑𝑑𝑑 = 𝐹𝐹 𝑑𝑑𝑑𝑑 and integrating we obtain the total impulse 𝐼𝐼: 𝐼𝐼 ≡ 𝐹𝐹 𝑑𝑑𝑑𝑑 = ∆𝑝𝑝  © NASA ROCKET ENGINES  Rockets and Space Exploration Rocket engines are the only current way to go into space. They have severe disadvantages: – Mostly non-reusable launchers (Exception: Falcon 9 and derivates, but more will join in the short term), – Expensive, – Modest performances, – Relatively unreliable (maximum reliability of around 99%), – Extreme complexity, – Expensive ground facilities and services. These disadvantages have an important effect on the engineering of satellites.  Going to London (shopping weekend) Let’s go buy some books in London. Conditions: – The plane will be boarded by, at most, seven people, including the crew. – The brand-new Airbus 350-900 will be trashed upon return to Barcelona (cost: 317.5 M€). We pay for it. – Chance of surviving the trip: around 99%. – Travel insurance: 25% of the total travel cost. © Lufthansa Who is coming?  Impacts on spacecraft design Due to the remarkable limitations of modern rockets, spacecraft design is jeopardised in several ways. Satellites must be very resilient, as there are no repair services in space. Then a. Parts and devices must be space-qualified (often expensive, power-hungry, massive, dated). De-rating to avoid systems overload. b. No mass production. c. Redundancies (hardware and functional). d. Extended missions and low reposition rate. e. Extensive ground testing. f. Long development times. g. Expensive insurance.  The thrust (1) Rockets transform stored energy (usually chemical) into kinetic energy. Newton’s second principle states 𝑑𝑑𝑝𝑝⃗ 𝐹𝐹⃗ = 𝑑𝑑𝑑𝑑  The thrust (2) If we take into account the forces caused by internal and external pressures we can obtain the thrust as 𝐹𝐹⃗ = − ∫𝐴𝐴 𝑃𝑃e 𝑑𝑑𝐴𝐴⃗ − ∫𝐴𝐴 𝑃𝑃𝑃𝑃𝐴𝐴⃗ + 𝐹𝐹⃗b e  The thrust (3) We can write it as 𝑣𝑣e 𝐹𝐹 = − 𝑃𝑃e − 𝑃𝑃a 𝐴𝐴e − ∫𝐴𝐴 𝑃𝑃 − 𝑃𝑃a 𝑑𝑑 𝐴𝐴⃗ + 𝐹𝐹⃗b ⃗ 𝑣𝑣e If we take the time derivative of linear momentum, we get 𝑑𝑑 𝑝𝑝⃗ 𝑚𝑚 + ∆𝑚𝑚 𝑣𝑣⃗ + ∆𝑣𝑣⃗ + −∆𝑚𝑚 𝑣𝑣⃗ + 𝑣𝑣⃗e − 𝑚𝑚𝑣𝑣⃗ = lim = 𝑚𝑚𝑣𝑣⃗̇ − 𝑚𝑚̇ 𝑣𝑣⃗e 𝑑𝑑𝑑𝑑 ∆𝑡𝑡→0 ∆𝑡𝑡 And this allows us to write 𝑣𝑣 𝑚𝑚𝑣𝑣⃗̇ = − ∫𝐴𝐴 𝑃𝑃 − 𝑃𝑃a 𝑑𝑑 𝐴𝐴⃗ − 𝑃𝑃e − 𝑃𝑃a 𝐴𝐴e e + 𝐹𝐹⃗b + 𝑚𝑚̇ 𝑣𝑣⃗e 𝑣𝑣e  The thrust (4) We define the aerodynamic force as 𝐹𝐹⃗a ≡ − 𝑃𝑃 − 𝑃𝑃a 𝑑𝑑 𝐴𝐴⃗ 𝐴𝐴 And then 𝑚𝑚𝑣𝑣⃗̇ = 𝐹𝐹⃗a + 𝐹𝐹⃗b + 𝑇𝑇 Where 𝑇𝑇 is the thrust 𝑣𝑣⃗e 𝑇𝑇 = − 𝑃𝑃e − 𝑃𝑃a 𝐴𝐴e + 𝑚𝑚̇ 𝑣𝑣⃗e 𝑣𝑣e  The thrust (5) 𝑣𝑣e 𝑇𝑇 = − 𝑃𝑃e − 𝑃𝑃a 𝐴𝐴e + 𝑚𝑚̇ 𝑣𝑣⃗e momentum thrust 𝑣𝑣e pressure thrust Momentum thrust is always larger than pressure thrust A rocket gets maximum thrust when pressure thrust is zero (matched nozzle)  The thrust (6) The effective ejection velocity is defined as 𝐴𝐴e 𝑣𝑣⃗e 𝐶𝐶⃗ = − 𝑃𝑃e − 𝑃𝑃a + 𝑣𝑣⃗e 𝑚𝑚̇ 𝑣𝑣e and then the equation of motion reads 𝑚𝑚𝑣𝑣⃗̇ = 𝐹𝐹⃗a + 𝐹𝐹⃗b + 𝑚𝑚̇ 𝐶𝐶⃗  Specific Impulse (1) The specific impulse is a measure of the energetic efficiency of a propulsion system (fuel + rocket engine). It is defined as ∫ 𝑇𝑇 𝑑𝑑𝑑𝑑 𝐼𝐼 𝐼𝐼sp = = 𝑔𝑔0 ∫ 𝑚𝑚̇ 𝑑𝑑𝑑𝑑 𝑔𝑔0 ∫ 𝑚𝑚̇ 𝑑𝑑𝑑𝑑 where 𝑔𝑔0 is gravity’s acceleration at sea level: 𝑔𝑔0 = 9.80665 m/s2. We can also use that 𝑇𝑇 ≈ 𝑚𝑚𝑣𝑣 ̇ e , and then 𝑚𝑚𝑣𝑣 ̇ e 𝑣𝑣e 𝐼𝐼sp ≈ = 𝑚𝑚𝑔𝑔 ̇ 0 𝑔𝑔0  Specific Impulse (2) The specific impulse is measured ins seconds or, better still, in N/(N/s). If we consider that the rocket works at constant 𝑇𝑇 and 𝑚𝑚,̇ we can write 𝑚𝑚̇ 𝐶𝐶 ∆𝑡𝑡 𝐶𝐶 𝐼𝐼sp = = 𝑔𝑔0 𝑚𝑚̇ ∆𝑡𝑡 𝑔𝑔0  Specific Impulse (3) Propellant 𝐶𝐶 (km/s) 𝐼𝐼sp (s) 𝑇𝑇 (N) Example Cold gas 0,5 – 0,7 50 – 70 0,1 – 200 N2, NH3 Solid 2,5 – 2,9 250 – 300 50 – 5×106 Al + NH4ClO4 Liquid 2,0 – 2,2 200 – 225 0,1 – 1000 N2H4, H2O2 (monopropellant) Liquid 2,9 – 4,4 300 – 450 50 – 5×106 H2 + O2 (bipropellant) Electrostatic 12 - 50 1200 – 5000 10-3 – 10 Ar, Xe, Kr, Hg  Rocket equation (1) Assuming that there are no external forces and that ejected gas expands up to 𝑃𝑃 = 0 Pa in vacuum, the equation of motion reduce to 𝑚𝑚f 𝑣𝑣 𝑑𝑑𝑑𝑑 −𝑣𝑣e = 𝑑𝑑𝑑𝑑 𝑚𝑚i 𝑚𝑚 0 where it has been assumed zero initial velocity. This equation can be integrated to yield Tsiolkovsky’s or rocket’s equation: 𝑚𝑚i 𝑣𝑣 = 𝑣𝑣e ln 𝑚𝑚f  Rocket equation (2) We can then write 𝑚𝑚i ∆𝑣𝑣 = 𝑔𝑔0 𝐼𝐼sp ln 𝑚𝑚f where we have used that 𝑣𝑣e = 𝑔𝑔0 𝐼𝐼sp 𝑚𝑚 The quantity i is called mass-ratio, and is one of the most important metrics to 𝑚𝑚f assess the performance of a launcher.  Rocket equation (3) The speed required to enter in low Earth orbit is, approximately, 8 km/s. For a system with a (very good) specific impulse of 450 s, the required mass ratio is 𝑚𝑚i ≈ 6.14 𝑚𝑚f but taking into account gravitational and atmospheric losses (that amount to about 2000–3000 m/s) we obtain 𝑚𝑚i ≈ 9.66 𝑚𝑚f  Embedded problem 2 A single-stage rocket must provide a total ∆𝑣𝑣 of 2.38 km/s to escape from the Moon. From a Lunar polar base, a team is preparing an LH/LOX rocket with a specific impulse of 415 seconds (in vacuum; and measured in the company premises near Samara, Russia). Determine the mass ratio required to escape our satellite. Considering that the radius of the Moon is 1737.1 km, and its rotational period is 27.322 days, do you think that locating the launch base near the Lunar equator is worth the effort?  Embedded problem 3 Assuming that the thrust of Saturn V’s diverse stages was approximately constant during their respective ignitions (each generating different thrust levels), why was there these increases in the accelerations provided by the launcher’s stages? © NASA  Recap: Rocket Engines There are several figures of merit for rocket’s performances: Thrust: 𝑇𝑇 = 𝑚𝑚𝑣𝑣 ̇ e + 𝑃𝑃e − 𝑃𝑃a 𝐴𝐴e Specific 𝑇𝑇 𝐶𝐶 𝐼𝐼sp = = impulse: 𝑚𝑚𝑔𝑔 ̇ 0 𝑔𝑔0 𝑚𝑚i DeltaV: ∆𝑣𝑣 = 𝑔𝑔0 𝐼𝐼sp ln 𝑚𝑚f Most rockets are expendable, expensive and relatively unsafe. This results in serious limitations in spacecraft design.  © Unknown ROCKET TECHNOLOGIES  General layout of a rocket engine We will discuss several parts of the Thrust rocket engine: 𝑣𝑣e 1. Combustion chamber 2. Nozzle 3. Tanks Combustion Nozzle chamber 4. Propellants (solid vs liquid)  Combustion chamber (1) Combustion chamber is the place where the propellants are burned. Both, structural and thermal control issues are severe problems. Pressure in the combustion chamber is high because: - it fosters a more compact design for the engine, - allows a larger flexibility for the nozzle’s aperture ratio, - gases are ejected at a higher velocity.  Combustion chamber (2) Chamber’s pressure is high even if: – Structural and thermal control problems are more severe, – They are prone to develop combustion instabilities, – Turbopumps, if required, must be more powerful, – Fuel leaks are potentially more detrimental. © NASM  Area-velocity relation (1) The continuity equation for a compressible fluid in a duct can be written as 𝑚𝑚̇ = 𝜌𝜌 𝐴𝐴 𝑣𝑣 = constant Then, making the logarithmic derivative 𝑑𝑑𝑑𝑑 𝑑𝑑𝜌𝜌 𝑑𝑑𝑑𝑑 + + =0 𝐴𝐴 𝜌𝜌 𝑣𝑣 If viscosity and external forces are neglected (both reasonable assumptions), Euler’s equation is 𝜌𝜌 𝜕𝜕𝑡𝑡 𝑣𝑣⃗ + 𝑣𝑣⃗ ∇𝑣𝑣⃗ + ∇𝑃𝑃 = 0 Finally, we deduce that 𝑑𝑑𝑑𝑑 2 𝑑𝑑𝑑𝑑 = Ma − 1 𝐴𝐴 𝑣𝑣  Area-velocity relation (2) For Ma < 1, a section reduction increases the gas speed. In the nozzle’s throat, we get Ma = 1. After the throat, the nozzle expands to obtain supersonic speeds (Ma > 1). 𝑑𝑑𝑑𝑑 2 𝑑𝑑𝑑𝑑 = Ma − 1 𝐴𝐴 𝑣𝑣  Nozzle efficiency Nozzles are designed to be optimal at a given height. We define optimal aperture ratio at height h as that at which 𝑃𝑃a = 𝑃𝑃e. Then, in this situation: 𝑇𝑇 = 𝑚𝑚̇ 𝑉𝑉e 1. Optimal aperture 2. Aperture ratio too high: flux separation 3. Aperture ratio too low: turbulence  Adaptive nozzles Currently, there are no commercial launchers that use adaptive nozzles to allow an optimal aperture ratio at several heights. There are, however, some that use spikes or aerospikes (this is the case of Barcelona-based Pangea: https://www.pangeaaerospace.com/). Adaptive nozzles are not yet used due to its complexity, and because the extra mass causes a penalty.  Nozzle optimization There must always be a compromise with nozzle optimization: - Nozzles designed for atmospheric ascent can only be optimal at a given height (that is, for a given external pressure), - Nozzles designed for its use in space (P∞=0) would require an infinite aperture ration. In this case, the limitations are nozzle’s mass and volume, - In lower stages, available volume has a significant effect on nozzle’s aperture ratios.  Fuel tanks (1) Tanks devoted to store liquid fuels use to be spherical or cylindrical (often, this last class uses to have semi-spherical ends). Tank’s mass will be determined by the stresses it must endure. These stresses can be given as the load per unit surface. If the wall has a thickness t and an internal radius r, an internal pressure of P results in an stress given by 𝑃𝑃𝑃𝑃𝑟𝑟 2 𝑃𝑃𝑃𝑃𝑃𝑃 𝜎𝜎 = = 2𝜋𝜋𝜋𝜋𝜋𝜋 2𝜋𝜋𝜋𝜋  Fuel tanks (2) Therefore, the tank’s mass is 4𝜋𝜋 3 𝑀𝑀 = 𝜌𝜌 𝑟𝑟 + 𝑡𝑡 − 𝑟𝑟 3 3  Fuel tanks (3) In space, surface tension becomes one of the dominant phenomena in fluid dynamics: https://www.youtube.com/watch?v=bKk_7NIKY3Y To avoid fluid bubbles hampering fuel flow, several methods are employed to force the flow (most of them based on elastic membranes or on the pressure of an inert gas). Simple, cheap method. On the other hand, efficiency changes with gas pressure if not controlled.  Recap: Rocket parts There are diverse components in a rocket engine: 1. Combustion chamber: high pressure and temperature to provide enough energy to the exhaust gas. Gas bulk velocity is very subsonic. 2. Convergent-Divergent Nozzle: increases the bulk velocity of the exhaust gases to supersonic velocity. 3. Propellant tanks: reservoirs of propellants. They face specific problems (thermal control, internal pressure, microgravity issues, corrosion, mass minimisation, among others).  © Roscosmos CLASSES OF ROCKETS  Classes of rockets Basically, there are two kind of rockets, although several others have been designed: – Solid rockets – Liquid rockets (mono- and bi-propellant) – hybrid – Electric propulsion – nuclear – other (cold gas, solar or laser sails, …) Their characteristics make them suitable for different tasks.  Solid motors (1) Fuel is Al powder, and oxidizer is ammonia perchlorate (NH4ClO4); fuel and oxidizer are bound by means of a binder (often HTPB, Hydroxi-Terminated Poly- Butadiene). The performance is intermediate, but the storage easiness makes them attractive for certain tasks. The fuel is (almost) always drilled, in such a way that it is possible to control the thrust and its time-evolution with the geometry of the perforation. Once the grain is cast, thrust is fixed. The perforation is called port. Its main disadvantage is that solid motors cannot be stopped once ignited.  Solid motors (2) In a solid, the combustion chamber is the port. And, obviously, the chamber’s volume increases with time. There is always a remaining of unburnt fuel, called sliver. In a well-designed rocket, the sliver mass is less than a 0.5% of the total rocket’s fuel mass. Solid engines use ablative nozzles for thermal control. Nozzle’s throat is subjected to erosive processes by alumina grains. To avoid thermal control problems, the grain is enshrouded by a layer of pure binder. Then, combustion is extinguished at a safe distance of the rocket structure.  Solid motors (3) The geometry of the port determines the total thrust and its time evolution (regressive, neutral, progressive rockets). If more than one ignition is required, the fuel can be separated in two or more grains by means of layers of pure binder; when the burning front gets to these layers, the burning extinguishes. Subsequent re-ignitions will proceed –when required– by means of the standard ignition procedure. The exhaust gases tend to by dark grey due to the formation of soot and the ejection of small particles of propellant. Solid rocket grain with star-like port  Combustion front dynamics Considering the previous expressions, the weight of fuel burned per unit time is: 𝑤𝑤̇ 𝑝𝑝 = 𝐴𝐴𝑏𝑏 𝜌𝜌 𝑟𝑟 being Ab the area of the port, ρ fuel’s density and r the advance speed of the burning front. Hence, the thrust will be: 𝑇𝑇 = 𝑤𝑤̇ 𝑝𝑝 𝐼𝐼𝑠𝑠𝑠𝑠 = 𝐴𝐴𝑏𝑏 𝜌𝜌 𝑟𝑟 𝐼𝐼𝑠𝑠𝑠𝑠  Combustion front dynamics The pressure in the combustion chamber complies that: 1 𝐴𝐴𝑏𝑏 1−𝑛𝑛 𝑃𝑃𝑐𝑐 ∝ with 0.2 ≤ 𝑛𝑛 ≤ 0.8 𝐴𝐴𝑡𝑡 which shows the risk of explosion for 𝑛𝑛 = 0.8, as the exponent is 5.  Solid motor cons Solid propulsion engines have several risks and disadvantages: – Ageing of the fuel or a production defect can give rise to small cracks in the grain. Their effect could be catastrophic, because they increase the burning surface with a negligible volume available. – They are impossible to stop. – Thrust is fixed by the geometry of the port and the characteristics of the fuel. – Their specific impulse is not very high Vulcan launch anomaly: https://www.youtube.com/watch?v=ZPztD5zwgYY.  Liquid engines (1) They give a higher 𝐼𝐼sp as compared with solid motors, and also allow changes in thrust, as well as stops and re-ignitions. Liquid engines are far more complex than solid motors. There exists several combinations of fuel/oxidizer; one giving a very high 𝐼𝐼sp is the combination LOX/LH (H2/O2). Kerolox (kerosene and liquid oxygen) and methalox (methane and liquid oxygen) are also widely used.  Liquid engines (2) There are two ways to feed fuel inside the combustion chamber: 1. Inert gas pressure (often nitrogen or helium): a tank at very high pressure feeds gas to the fuel tanks, that are then forced into the combustion chamber This technique is rather dated. © Tacca, Rachov & Lentini © G. P. Sutton  Liquid engines (3) 2. Propellants can be forced into the combustion chamber by means of turbopumps driven by a hot gas generator that burns small quantities of propellants. To start the turbopumps, a separate gas generator (often a small solid propellant rocket) is used. There can be open cycle and closed cycle systems. In open cycle engines, the exhaust of the turbopump is injected in the nozzle, while in closed cycle systems it is injected (at high pressure) in the combustion chamber. Open cycle systems are simpler, closed cycle systems slightly improve the performance.  Vulcain Engine Open cycle System Hydrolox (LOX/LH) 𝐼𝐼sp= 431 s (vac) 𝐼𝐼sp= 326 s (sea level) 𝑇𝑇= 1,075.00 kN Burning time: 600 s 𝑇𝑇comb= 3250℃ 𝑃𝑃 ≥ 100 bar Combustion efficiency higher than 99% Liquid engines: classification (1) Storable propellants: hydrazine, hydrogen peroxide, alcohol, kerosene. – Stable at moderate temperature. – Often toxic and/or corrosive. Cryogenic propellants: liquid oxygen, liquid hydrogen. – Stable at very low temperature. – Non-toxic, escape gases non-hazardous. Liquid engines: classification (2) Monopropellant: based on the catalysed dissociation of the reaction fluid. Usually, hydrazine (N2H4) or hydrogen peroxide (H2O2). Bipropellant: they combine a fuel and an oxydizer: – Liquid oxygen and liquid hydrogen (LOX/LH). – Liquid oxygen and kerosene (KeroLox). Tripropellant: a bipropellant system with the addition of catalysts.  Hypergolic propellants Some compounds spontaneously react when in contact, which allows to avoid ignition systems that can fail. A widely used combination is nitrogen tetroxide (N2O4) and hydrazine (N2H4). The resulting exhaust is quite toxic, which results in wide avoidance zones for operators during the launch. The fumes are orange-hued, as is N2O4. This combination is used in several members of the Long March family, and was used in Titan and LEM, among other examples. © CGWC Launch of a Chinese Long March 3 rocket. The orange cloud is composed of toxic gases, which results in complex operational procedures.  Wet dress rehearsal In liquid engines, a wet dress rehearsal is a fuel loading test, without the ignition of the rocket. This kind of test is used to proof that loading is flawless, that there are no leaks, that tanks withstand the stress (mechanical as well as thermal)…  Static tests Test of a cryogenic liquid engine (LH+LOX) https://youtu.be/08Gv7qDxgUE Failures of static tests: https://www.youtube.com/watch?v=1mTCxpGSIbI  Static tests PLD Space is using a static test stand in Teruel’s airport. In Catalunya, the Generalitat has similar plans with Lleida’s airport. So far, Cosmic Research (a student association in the Terrassa Campus) have performed solid engine tests. Static tests © PLD Space (all)  Regenerative cooling Invented by Robert H. Goddard. It consists in the circulation of one of the fuels through a cooling circuit (skirt) that enshrouds the external part of the nozzle: – Excellent thermal control, – Fuel pre-heating, – Only usable in bipropellant systems.  Monopropellant systems Based on the dissociation of high-energy molecules. Hydrazine flowing over a catalyst bed (often made of iridium platinum microspheres): 2 N2H4 N2+ 2 NH3+ H2 The resulting products are (rather) innocuous. Astrium hydrazine vernier Used for satellite attitude control. Characteristics: – Thrust: 7.9 – 24.9 N – Specific impulse: 230 s – Total impulse: 517 kN·s – Mass: 0.395 kg Hydrogen peroxide as a propellant  Hybrid rockets (1) In this case, the fuel is a solid and the oxydiser is a liquid or a gas. The have several advantages: reliable, throttle can be controlled, they can be stopped and re-started, are storable and (rather) clean. They combine many advantages of solid and liquid engines. Nevertheless, its use has been quite limited until SpaceShip One and Two.  Other propulsion methods There are other kinds of rockets, but we will not discuss them: 1. Electromagnetic engines, 2. Nuclear rockets, 3. Solar sails and laser sails, 4. Magnetic sails  Recap: Rocket types Several technologies are used for rockets: 1. Solid motors: long storage times feasible, intermediate-high performance, comparatively simple, useful for launchers. 2. Liquid bipropellant engines: non-storable propellants or storables with relevant safety issues, high performance but very complex, useful for launchers. 3. Liquid monopropellant engines: storable propellants, low thrust and low to intermediate specific impulse, useful for attitude control. 4. Cold gas: very low efficiency and thrusts, only useful for attitude control and special applications (such as the Manned Maneuvering Unit).  © CGWC LAUNCHERS  © ISRO Fairing Anatomy of a launcher We will consider the following parts for a launcher: 1. Fairing, Stages 2. Stages and boosters, 3. Rocket engines (already covered). Avionics and software remain out of scope for this topic. Boosters  The shroud A rocket’s shroud is designed to protect satellites from the air flow at high Mach number. This flow would damage the satellites due to aerodynamic forces and aerothermal heating. As rockets leave the densest layers of the atmosphere in a few minutes, shrouds are rapidly ejected (as, being on top of the rocket, its mass would affect all rocket phases). Fairing separation: https://x.com/torybruno/status/1843379189803839678. Atlas V shroud (satellites: Magnetospheric Multidimensional Explorers). The shroud (2) The shroud (3) Aerothermal Flux (1) Aerothermal Flux (2) Aerothermal Flux (3) Aerothermal Flux (4)  Stage rockets (1) We have seen that in order to reach orbital velocity a single phase rocket would require an unrealistic mass ratio. Tsiolkovsky noticed that putting rockets on top of rockets it would be possible to reach Earth orbit. Given the central role of structural mass, the purpose of step rockets is to eject parts of the structure in different phases of the launch http://www.space.com/27855-orion-spacecraft-flight-full- coverage.html#ooid=FpYmI0cjpaQkIjB2-JGnDsGBqN-ZWnqn  Stage rockets (2) It is possible to generalize the previous definitions of structural mass and payload ratio for each phase: 𝑚𝑚𝑠𝑠,𝑘𝑘 𝜖𝜖𝑘𝑘 = 𝑚𝑚𝑠𝑠,𝑘𝑘 + 𝑚𝑚𝑝𝑝,𝑘𝑘 𝑚𝑚0,𝑘𝑘+1 𝜋𝜋𝑘𝑘 = 𝑚𝑚0,𝑘𝑘 𝑚𝑚𝑃𝑃𝑃𝑃 𝜋𝜋𝑃𝑃𝑃𝑃 = 𝑚𝑚0,1  Stage rockets (3) With these definitions, each stage can be treated as a single-stage rocket Then, the final velocity will be the sum of final velocity of each individual stage: 𝑁𝑁 𝑉𝑉f = − 𝑉𝑉e,𝑘𝑘 ln 𝜖𝜖𝑘𝑘 + 1 − 𝜖𝜖𝑘𝑘 𝜋𝜋𝑘𝑘 𝑘𝑘=1 The total mass-ratio will be 𝑵𝑵 𝜋𝜋PL = 𝜋𝜋𝑘𝑘 𝑘𝑘=1  Stage rockets (4) The condition to obtain large final velocities (even larger than the gas ejection velocity) is to have a small payload mass-ratio. This can be obtained through several stages of intermediate payload mass-ratio. Performance increases with the number of stages, but after 3 or 4 stages the improvement is modest, and the complexity (and cost!) grows very fast.  Stage rockets (5) In several cases, there are boosters around the first stage. Then, we talk about a zero-th stage, and get 𝑇𝑇tot = 𝑚𝑚̇ 0 𝑉𝑉e,0 + 𝑚𝑚̇ 1 𝑉𝑉e,1 = 𝑚𝑚̇ tot 𝑉𝑉 e,0 where 𝑉𝑉 e,0 is an averaged velocity for the stages that work in parallel: 𝑚𝑚̇ 0 𝑉𝑉e,0 + 𝑚𝑚̇ 1 𝑉𝑉e,1 𝑉𝑉 e,0 = 𝑚𝑚̇ 0 + 𝑚𝑚̇ 1  Stage rockets (6) The effective values for the structural and payload mass-ratios are 𝑚𝑚s,0 + 𝑚𝑚s,1 𝜖𝜖0 = 𝑚𝑚s,0 + 𝑚𝑚s,1 + 𝑚𝑚p,0 + 𝑚𝑚ip,1 𝑚𝑚01 − 𝑚𝑚ip,1 𝜋𝜋0 = 𝑚𝑚00 where we assume that only part of the fuel is burnt while the stages work simultaneously (𝑚𝑚ip,1 ).  Stage rockets (7) For the first stage, we have 𝑚𝑚s,1 𝜖𝜖1 = 𝑚𝑚s,1 + 𝑚𝑚p,1 − 𝑚𝑚ip,1 𝑚𝑚02 𝜋𝜋1 = 𝑚𝑚02 + 𝑚𝑚s,1 + 𝑚𝑚p,1 − 𝑚𝑚ip,1  Stage rockets (8) With these definitions, the final velocity is 𝑁𝑁 𝑉𝑉f = −𝑉𝑉 e,0 ln 𝜖𝜖0 + 1 − 𝜖𝜖0 𝜋𝜋0 − 𝑉𝑉e,𝑘𝑘 ln 𝜖𝜖𝑘𝑘 + 1 − 𝜖𝜖𝑘𝑘 𝜋𝜋𝑘𝑘 𝑘𝑘=1  Multi-stage launcher optimization (1) The final velocity of the rocket is given as 𝑁𝑁 𝑁𝑁 𝑚𝑚0,𝑘𝑘 𝑉𝑉f = 𝑉𝑉e,𝑘𝑘 ln = − 𝑉𝑉e,𝑘𝑘 ln 𝜖𝜖𝑘𝑘 + 1 − 𝜖𝜖𝑘𝑘 𝜋𝜋𝑘𝑘 𝑚𝑚f,𝑘𝑘 𝑘𝑘=1 𝑘𝑘=1 we can maximize this velocity by modifying 𝑉𝑉e,𝑘𝑘 , 𝜖𝜖𝑘𝑘 , and/or 𝜋𝜋𝑘𝑘. Nevertheless, the ejection velocity and the structural mass-ratio are determined by the available technologies. Then, payload mass is the most relevant parameter. The most practical mathematical tool for this purpose are the indeterminate Lagrange multipliers.  Multi-stage launcher optimization (2) The problem of conditioned optimisation reduces then to maximize the payload mass-ratio 𝑁𝑁 𝜋𝜋PL = 𝜋𝜋𝑘𝑘 𝑘𝑘=1 If we take the logarithm of the previous expression, the product turns into an addition, and we can apply the method of Lagrange undetermined multipliers: 𝑁𝑁 𝑉𝑉f ln 𝜋𝜋PL = ln 𝜋𝜋𝑘𝑘 + 𝜆𝜆 + 𝑉𝑉e,𝑘𝑘 ln 𝜖𝜖𝑘𝑘 + 1 − 𝜖𝜖𝑘𝑘 𝜋𝜋𝑘𝑘 𝑁𝑁 𝑘𝑘=1  Multi-stage launcher optimization (3) After some algebra, we obtain 𝑒𝑒 −𝛽𝛽 𝜆𝜆 = 𝑉𝑉e 𝜖𝜖 − 𝑒𝑒 −𝛽𝛽 𝑉𝑉f 𝛽𝛽 = 𝑁𝑁𝑉𝑉e where we have assumed that all stages eject gas at the same velocity and have the same structural mass-ratio. The optimal payload mass-ratio for each stage is then 𝑒𝑒 −𝛽𝛽 − 𝜖𝜖 𝜋𝜋𝑘𝑘 = 1 − 𝜖𝜖  Multi-stage launcher optimization (4) Empirically, we find that: – Stages with larger specific impulse must be over the ones with lesser 𝐼𝐼sp. – The stages with higher 𝐼𝐼sp must give larger ∆𝑣𝑣. – Each successive stage must have less mass than the preceding ones – Similar stages must give similar ∆𝑣𝑣.  Figures of merit The most relevant metric for a launcher is the mass into orbit: PSLV (ISRO): – Payload to LEO: 3800 kg – Payload to SSO: 1750 kg – Payload to GTO: 1425 kg Besides, the accuracy of the orbit injection is also critical (semimajor axis, eccentricity and inclination). For Ariane 5, we had ∆𝑎𝑎 = 2.5 km, ∆𝑒𝑒 = 3.5 × 10−4 , ∆𝑖𝑖 = 0.04°, and ∆Ω = 0.03°.  Figures of merit The launcher performance can be assessed by the 𝐶𝐶3 parameter. This is an astrodynamical parameter derived from the vis-viva equation: 𝑣𝑣 2 𝜇𝜇 1 𝜀𝜀 = − = constant = 𝐶𝐶3. 2 𝑟𝑟 2 So, the 𝐶𝐶3 is the square of the excess velocity in hyperbolic orbits: 2 𝐶𝐶3 = 𝑣𝑣∞ , and is related to the energy that must be provided by a launcher to a spacecraft to get to a specific destination.  © Hergé LAUNCHER DYNAMICS  Launcher Dynamics During the ascent trajectory, the rocket is subjected to several forces: – Thrust – Weight – Lift – Drag Thrust Weight Drag Lift  Equations of Motion 𝑑𝑑𝑑𝑑 𝐹𝐹 cos 𝛼𝛼 − 𝐷𝐷 𝛼𝛼 determines the launcher’s 𝑑𝑑𝑑𝑑 = 𝑚𝑚 − 𝑔𝑔 sin 𝛾𝛾 ascending trajectory and is a control 𝑑𝑑𝑑𝑑 𝐹𝐹 sin 𝛼𝛼 + 𝐿𝐿 𝑉𝑉 2 parameter, as is the case with 𝑉𝑉 = − 𝑔𝑔 − cos 𝛾𝛾 𝑑𝑑𝑑𝑑 𝑚𝑚 𝑟𝑟 𝑑𝑑𝑑𝑑/𝑑𝑑𝑑𝑑. 𝑑𝑑𝑑𝑑 𝑅𝑅 = 𝑉𝑉 cos 𝛾𝛾 𝑑𝑑𝑑𝑑 𝑟𝑟 𝛾𝛾 is the angle between the thrust (𝑇𝑇) 𝑑𝑑𝑑𝑑 𝑑𝑑𝑑 = = 𝑉𝑉 sin 𝛾𝛾 and the velocity vector (𝑣𝑣). 𝑑𝑑𝑑𝑑 𝑑𝑑𝑑𝑑 𝑑𝑑𝑑𝑑 = −𝑚𝑚̇ 𝑡𝑡 We have assumed that the Earth 𝑑𝑑𝑑𝑑 does not rotate. 1 2 𝐿𝐿 = 𝜌𝜌𝑉𝑉 𝑆𝑆𝐶𝐶𝐿𝐿 2 1 𝐷𝐷 = 𝜌𝜌𝑉𝑉 2 𝑆𝑆𝐶𝐶𝐷𝐷 2 2 𝑅𝑅⊕ 𝑔𝑔 = 𝑔𝑔0 𝑟𝑟 2 𝛼𝛼 = 𝛼𝛼 𝑡𝑡  © Fritz Lang & UFA MODERN LAUNCHERS  Modern launchers Modern launchers have improved their reliability, but the cost continues being too high for affordable access to space. In the XXI century, there is an increasing trend towards more and more private actors in the space arena, thus reducing the share of public institutions. Nevertheless, the main goal of these corporation is profit. Whether this is good, or evil, or both or none is still a matter of (heated) debate.  Falcon 9 (SpaceX) Mass to LEO is 22,800 kg. First stage and fairing are recoverable to reduce launch costs. The cost per launch is 50 million USD (62 previous to reusability). Propellants are cooled (O2 to 66.5 K and RP1 to 266.5 K) to increase their density and fit more mass in the reservoirs. This strategy is called “subcooling”. Success rate: 230/232 (but one rocket exploded in the launchpad during countdown). Landings: 206/211. © SpaceX  Falcon 9 (SpaceX) First stage 9 Merlin engines Thrust (sl) 7,607 kN Thrust (vac) 8,227 kN Isp (sl) 282 s Isp (vac) 311 s Burn time 162 s Propellants LOX+RP-1  Falcon 9 (SpaceX) Second stage 1 Merlin engine Thrust (vac) 934 kN Isp (vac) 348 s Burn time 397 s Propellants LOX+RP-1 © SpaceX © SpaceX  Falcon Heavy (SpaceX) It is a modification of the Falcon 9: it has three Falcon 9 first stages running in parallel (one is the first stage, the other two are boosters). Its mass to LEO is 63,800 kg; only the Saturn V and the Energia (both retired) had larger payload masses. Success rate: 8/8  Falcon Heavy (SpaceX) First stage 27 Merlin engines Thrust (sl) 7,607 kN Thrust (vac) 8,227 kN Isp (sl) 282 s Isp (vac) 311 s Burn time 162 s Propellants LOX+RP-1  Falcon Heavy (SpaceX) Second stage 1 Merlin engine Thrust (vac) 934 kN Isp (vac) 348 s Burn time 397 s Propellants LOX+RP-1 © SpaceX © SpaceX © SpaceX © SpaceX © SpaceX © SpaceX © SpaceX © SpaceX © SpaceX  Soyuz 2 Boosters (four) 4 RD 177 engines Thrust (sl) 1,021.097 kN Isp (sl) 310 s Burn time 120 s Propellants LOX+RP-1  Soyuz 2 First stage 4 RD 118 engines Thrust (sl) 999.6 kN Isp (sl) 311 s Burn time 286 s Propellants LOX+RP-1  Soyuz 2 Second stage 1 RD-0124 engine Thrust (vac) 294 kN Isp (vac) 359 s Burn time 300 s Propellants LOX+RP-1  Soyuz 2 Fregate (optional third stage) 1 S5.92 engine Thrust (vac) 19.6 kN Isp (vac) 325 s Burn time 877 s Propellants N2O4+UDMH © Roscosmos © Roscosmos © Roscosmos © Roscosmos © Roscosmos © Roscosmos © Roscosmos © Roscosmos © Roscosmos © Roscosmos © Roscosmos  New Shepard New Shepard is a fully reusable sounding rocket able to transport the Blue Origin Capsule to an altitude in excess of 100 km (thus, beyond the von Kármán line and into space). The engine is the BE-3 that produces 490 kN during the ascent and 90 kN during the controlled descent. In vacuum it is able to generate 710 kN. Its propellants are LOX/LH and the specific impulse has not been disclosed. The cost of an astronaut seat has not been revealed either. © Blue Origin © Blue Origin © Blue Origin © Blue Origin © Blue Origin © Blue Origin  © China Great Wall Corporation LAUNCH BASES  Launch bases (1) Launch bases must comply with a series of requirements: – Be far away from inhabited regions, – Next to the sea, – As long as possible, be close to the equator (for launches to low inclination orbits) or in high latitudes (for polar orbits). Russia and China do not comply with any of these requirements (geographical + strategic issues). Launch bases (2) Launch bases (3)  El Arenosillo El Centro de Experimentación de El Arenosillo (CEDEA) is a launch base for sounding rockets located in Huelva. It was used from 1966 to 1994 for the launch of sounding rockets, and more recently for the test of UAVs, among other purposes. Sounding rockets INTA 110, INTA 250 and INTA 300 were tested there. In total, more than 550 rockets were launched in this period. Recently, it has reopened as a launch base for the tests of PLD space (and ESEIAAT’s Cosmic Research as well). Its launch azimuth is not very convenient, as to avoid populated areas launches must be retrograde, which results in a substantial loss of efficiency of the rocket. © Google maps Allowed launch azimuths © Delft University Launch of a small rocket in 2015 by members of Delft University (Netherlands). This launch marked the return of rocket activities to the base.  Dog leg manoeuvre A dog leg is performed to avoid overflying certain regions because of Density of population, International policy issues, safety  ISRO’s dogleg manoeuvre to avoid endangering the population of Sri Lanka. Shown is a polar launch from Sriharikota. © ISRO  Chinese launch bases Chinese launch bases are located well within the mainland, not close to the coast, to avoid the possibility of a foreign army getting them under control. Then, the launch path of Chinese rockets are over densely populated regions of the country, a situation that has been the cause of several accidents (some involving deaths). The launch bases in China and their distance to the sea are: Jiuquan SLC (2000 km), Taiyuan SLC, Xichang SLC , and Wenchang SLC (10 km). To avoid this kind of problems (among other reasons), China is also developing ships for sea launches. The launch of a Long March 2 to the Chinese space station resulted in the first rocket stage falling to the ground very close to a village. The orange fumes and the cloud expanding after the impact, very toxic, posed a serious hazard to local population. Space debris found on the ground can be very dangerous. Only specialized crews should get close to them. In 1996, a Long March 3B launched from Xichang SLC veered off course and crashed on a small village at 1.2 km of the base, killing at least six people (official figures, but other evaluations increase the tally to a few hundred).  © THE NEAR FUTURE  StarShip (Space X) StarShip is a super-ambitious plan to build a single stage to orbit with a huge load capacity. So far, only a few parts (engines and some elements of attitude determination and control, among others) have been tested. © SpaceX © SpaceX © SpaceX An early version of StarShip (left), and a cutaway of the orbital model stacked on top a Falcon Heavy (right). © Spinlaunch SPINLAUNCH  Spinlaunch: back to Verne? The basic idea is simple, but has to overcome several serious problems: a centrifuge accelerates the payload to a high velocity and ejects it. Instead of 8 km/s required for a LEO orbit (neglecting gravitational and, especially, atmospheric losses), which would result in the collision of the satellite with the atmosphere, the payload is accelerated to 2.2 km/s and ejected in a suborbital trajectory. Once outside the atmosphere, the payload is deployed and a rocket allows the satellite to achieve LEO. © George Melies  Spinlaunch: back to Verne? So far, the company has conducted tests up to Ma ≈ 6. The nose cone of the projectile is made of wolfram and acts as a mechanical shield and thermal sink (to avoid problems related to aerothermal heating). At the tip of the tether, centripetal acceleration amounts to 10,000 G.  Embedded problem Compute the centripetal acceleration to which payloads launched with SpinLaunch will be subjected. Recall that 𝑣𝑣 2 𝑎𝑎cen =. 𝑟𝑟  Single Stage to Orbit (SSTO) Up to now, space launchers require several stages to achieve orbital velocity. A reusable SSTO could imply a substantial reduction in launch costs (today at about ~10 000 USD per kg). Or perhaps not: the Space Shuttle was the most expensive launch system ever designed (Hangar Queen).  Single Stage to Orbit (SSTO) There are two alternatives for SSTO design: - Vertical take-off/Horizontal landing, - Vertical take-off/Vertical landing. Vertical take-off allows to get rid of the wings required to sustain the launcher (including propellants) at low velocities. Horizontal landing allows a reduction on the fuel for vertical landing and the extra safety of gliding. In the case of launch abort, wings would allow a glide path compatible with soft landing (examples: Skylon, X33). © Boeing  Single Stage to Orbit (SSTO) Vertical landing allows to save the weight of wings (which causes a larger penalty than the extra fuel mass, as the engines are necessary also for take- off). It does not allow gliding, but a parachute system could provide viable alternative for emergencies. Example: Delta-Clipper X.  Electromagnetic launchers (1) Electromagnetic launchers (also called railguns or mass drivers) accelerate masses to a very high velocity. They sport several advantages: – Complete reusability, – Low cost (in energy terms), – High availability.  Electromagnetic launchers (2) Main disadvantages are related with the high accelerations to which payloads are subjected (up to 15000 g), as well as the intense heat experienced by the payload (unless launched from vacuum). The Moon is an excellent location for these mass drivers.  © Virgin Orbit Launchers for Small Satellites  Launchers for SmallSats The properties of interest of modern launchers for small satellites are: 1. Cost 2. Launch environment 3. Accuracy of injection (semimajor axis, orbital inclination, right ascension of the ascending node) 4. Secondary payloads accepted? 5. Dead launch? 6. Range of on-site services: clean rooms, S/C checkout equipment, filling hardware…  Launchers for SmallSats There exist several launchers that are specially well suited for nano and picosatellites: Decommissioned ICBMs (mostly from Russia/FSU) New, low cost, low performance launchers Piggyback launches Shared small launchers (Vega) Costs should be smaller than 50,000 USD/kg to be competitive with piggyback launches. A higher cost might be acceptable if launch date and/or final orbit can be chosen by the user.  PSLV The Polar Satellite Launch Vehicle is ISRO’s most successful launcher: Payload to LEO: 3800 kg Payload to SSO: 1750 kg Payload to GTO: 1425 kg Rate of success: 0.94 (50/53; 2 failures, 1 partial failure). Cost: 18 million US dollars (130 crore rupees).  © ISRO PSLV Strap-on motors Second stage Fourth stage (PSLV-G and PSLV-XL) Liquid (UDMH+N2O4) Liquid (MMH+MON) First stage Third stage Solid (HTPB) Solid (HTPB) © ISRO  Vega VEGA is ESA’s newest launcher. Italy: 65%, France: 13%, smaller contributions by Spain, Belgium, Nederland, Switzerland and Sweden. Payload to polar orbit: 1430 kg. Payload to SSO: 1450 kg. Payload to LEO (1500×200, i=5.4°): 1936 kg. Rate of success: 0.88 (15/17; 2 failures). Cost: 35 million US dollars. © ESA © ESA Vega Solid stages Liquid stage Al+NH4PO4+HTPB UDMH+N2O4  PLD Space A Spanish company developing two small satellite launchers, Miura 1 and 5. Miura 1 (former Arion 1) is a single stage, reusable suborbital rocket – Payload: 200 kg, – Engine: kerosene + LOX, 30 kN, – First engine test: April 2018. Miura 5 (former Arion 2) will be a three stage rocket able to insert payloads in a 400–1200 km orbit – Payload: 50 to 150 kg, – Cost: 30–35 k$/kg (expected), – First engine test: 2020. © PLD Space © PLD Space  PLD Space Miura 1 is a suborbital rocket designed for short-duration experiments and tests. Being a liquid engine, the maximum longitudinal acceleration will be a 30% of the ones found in solid rockets. Start of microgravity (𝑎𝑎 < 10−4 g) at 80 km, apogee at 153 km; microgravity elapsed time 3.7–4.7 minutes. Soft splash in the sea. © PLD Space  Miura 1 This is a suborbital demonstrator, a precursor of the Miura 5 which will be able to inject cargo in LEO. Engine is based in liquid propulsion (KeroLOX). The test expectations were for an apogee of 80 km, but finally it was just 46 km (space begins, by legal convention, at a height of 100 km). Apparently, there was a last-minute decision to increase the launch angle to reduce the risks of the returning leg of the mission. The rocket was lost in the Atlantic. In future missions, PLD expects to recover it for reuse. © PLD  PLD Space PLD space was founded in Elche in 20 by Raul Torres and Raul Verdú to design and operate a small launcher for minisatellites. Their first (suborbital) rocket is Miura 1 (formerly, Arion, but the name was changed due to its similarity to Ariane). After tests with Miura 1, they will move to Miura 5, this time an actual launcher to LEO. Company’s motto: Opening space for everyone… Provided they have the cash (my contribution)  Teprel 1B Miura 1 uses a single TEPREL-1B engine designed and manufactured by PLD providing 32 kN of thrust. Propellants are LOX (oxidiser) and Jet-A1 (commercial kerosene). Regenerative cooling system. The nozzle can be tilted for active thrust vector control. © PLD Expected, but not achieved, flight path for the test mission of Miura 1.  PLD Space Miura 5 will be a launcher for payloads of up to 300 kg and will allow dedicated, piggyback and rideshare launches. Recoverable first stage (not yet tested). Kick stage option. Operational launches will be made from Kourou. Up to 15 launches per year (projected). © PLD Space Artist image of Miura 5  Miura 5 PLD is developing Miura 5, a reusable two- stage launcher: First stage: 20.37 m long, 1.8 m of diameter and 5 Teprel C engines (KeroLOX) giving 525 kN of trust. Second stage: 10.94 m long with one Teprel C engine providing 65 kN in vacuum. Up to 450 kg to SSO. Un to 900 kg to LEO. © PLD Space  PLD Space First launches will be made from El Arenosillo, near Huelva, and this implies a retrograde orbit: cos 𝑖𝑖 = sin 𝑎𝑎 cos 𝜆𝜆 where i is the orbital inclination, a the launch azimuth and λ the launch base latitude.  Pangea Aerospace Pangea Aerospace is a small firm located in Barcelona and which is designing a small, two-stage launcher with an aerospike. It will inject 150 kg at LEO. The aerospike, which is optimal at all heights, is 3D printed. Despite its mechanical complexity, the whole engine is printed in just two pieces (a standard design and construction would require over two hundred parts). The stages will be reusable, and will use methane and liquid oxygen (CH4 + LOX). Pangea Aerospace © Pangea Aerospace Left: cold flow test of the aerospike at DLR facilities. Right: drop test of first stage in Kiruna (Sweden). © Pangea Aerospace  Virgin orbit Virgin Orbit offers launches of small satellites (up to 300 kg at 500 km SSO, 500 kg at 230 km SSO) from an air launched vector. The advantages of air-delivered launchers are: 1. Initial altitude (37 000 feet, 11 300 m), 2. Lower air density, 3. Initial velocity of 800 km/h, 4. No limitations in orbital inclination, 5. Legal advantages.  Virgin orbit First mission on January 17th, 2021 was a success. An orbital test on May 25th, 2020 had failed (it only carried a payload mock-up). Cost: 12 million US dollars. Aircraft: modified Boeing 747 (Cosmic Girl). Rocket: LauncherOne; two liquid stages (RP-1 and LOX). © Virgin Orbit © Virgin Orbit  Cost of small launchers Are small launchers inexpensive enough for small missions? Small rockets are not necessarily (and are not in fact) more affordable in terms of cost per unit mass than large rockets, as private jets are not less expensive (in cost per passenger) than airliners. The «sweet spot» is probably in the range from Dnepr to PSLV (Sweeting, 2018). © Matt Groening and David X. Cohen PIGGYBACK LAUNCHES  Piggyback launches Piggyback launches are very cheap, even free They have many serious drawbacks: – It is not possible to choose orbit. – Prime contractor decides whether you go or not (Possible de-manifest near the launch date). – Specific cost is very high (up to 50 000 dollars per kilogram); this trend worsens with decreasing satellite mass. – Uneven availability. – Reduced interfaces and services. – “Dead” launch frequently required. © JAXA © Mitsubishi Heavy Industries Ltd Seven satellites piggy-backing on the 15th launch of Japan’s HII rocket (2003). © ISRO Launch of PSLV C37 with 104 satellites on board © ISRO  ASAP5 ASAP stands for Ariane Structure for Auxiliary Payloads. It flied with the Ariane 5. Specifically designed for micro (less than 120 kg) and minisatellites (less than 300 kg), but was able accommodate several nano and picosatellites when needed. Three main configurations: – 8 microsats: maximum mass 960 kg – 4 minisats: maximum mass 1200 kg – 2 mini + 6 micro: maximum mass 1320 kg The system is no longer operational. © ESA © Kosmotras  © Uderzo et Goscinny RIDESHARE LAUNCHES  Rideshare In a rideshare, several satellites share a single launcher, whose cost is divided (taking into account the mass) between the different users. Some expenses cannot be shared, as interface design, especial needs of satellites, insurance, etc. The distribution of satellites must locate the centre of mass of the payloads in the longitudinal axis of the launcher.  © SpaceX Transporter 1 mission (SpaceX) with 143 satellites on board. This is actually a multiple launch, not a piggyback launch.  © SpaceX A SpaceX rideshare starts at 1 million US dollars irrespective of the mass of the satellite. Cost starts increasing over 200 kg in a linear way.  © SpaceX Rideshare © ESA Rideshare structures to be offered for Vega, Vega-C and Ariane 6. DEPLOYERS  Small Satellite Mission Service Dispenser The SMSS dispenser will be tested in the second half of 2018. It will allow Vega and Vega C to launch several small satellites. Satellites in the mass range 1 kg to 400 kg will fit in the SMSS. © ESA  Last mile services Unless the satellite is the primary payload, it is rather unlikely that the final orbit will be close to the optimal one for the mission. Recently, some providers are offering space tugs (also known as Orbital Transfer Vehicles) to deliver the secondary or rideshare payloads to more suitable orbits. This reduces dependence on dedicated small launchers. Examples: Spaceflight’s Sherpa family, as well as Momentus (https:https://momentus.space/), and D-orbit (https://www.dorbit.space) among other companies. This is a fast-growing commercial niche.  Last mile services Satellite constellations will become the main customer to these services for orbital dissemination and management. A future possibility, albeit rather complex, is life-extension of satellites by means of in-orbit refuelling. Super-heavy rockets, launching hundreds of small satellites in a single rideshare, will increase the demand of last mile services. © Spaceflight Last mile services  Last mile services DETAIL CONTAINERIZED SATELLITE CLASS PAYLOAD TYPE 3U 6U 12U 50 kg 100 kg 150 kg 200 kg 300 kg 450 kg 750 kg 1000 kg LENGTH (cm) 34.05 34.05 34.05 80 100 100 100 125 200 300 350 HEIGHT/DIA (cm) 10 10 22.63 40 50 60 80 100 150 200 200 WIDTH (cm) 10 22.63 22.63 40 50 60 80 100 – – – MASS (kg) 5 10 20 50 100 150 200 300 450 750 1000 PRICE- LEO $145k $295k $595k $895k $975k $1,350k $1,350k $1,850k CALL CALL CALL PRICE- GTO $915k $1,400k $2,750k $4,600k $8,500k $9,800k $11,200k $14,000k CALL CALL CALL Prices in USD. *Prices may vary by launch provider, orbit and/or spacecraft technical requirements. © Spaceflight. Retrieved on December 11th, 2022.  Small satellite dispensers Launching a single small satellite in a dedicated rocket does not make sense (unless a microlauncher specifically designed to this goal is developed). Piggyback launches use to involve the ejection of several small satellites. If the satellites have exactly the same external geometry (plus a few requirements about mass distribution) they can be housed and ejected by a dispenser. © CalPoly  Small satellite dispensers The are several advantages: – in cost (per satellite) – allow multiple launches – in security issues (enhanced protection) – simplified launcher – satellite interface – well-tested injection system – beaurocracy (insurances, export licenses…)  Small satellite dispensers Currently existing dispensers: – P-POD (CalPoly) – ISIPOD (ISIS) – J-POD (JAXA) – SSPL (Space Shuttle Picosatellite Launcher, US DoD) – NLAS (Nanosatellite Launch Adapter System, NASA) – T-POD (University of Toronto Space Flight Laboratory) – X-POD (University of Toronto Space Flight Laboratory) © NASA Small satellite dispensers The US will put several P-PODs in every accessible launcher (Educational Launch of Nanosatellites, www.nasa.gov/mission_pages/smallsats/elana/). © Orbital Sciences Corporation Isipod mounted on the Antares.  PPOD PPOD stands for Poly Picosatellite Orbital Deployer, and was developed at the California Polytechnic University. Main goals – Protect LV and primary payload, – Protect CubeSats from launch environment, – Safe/reliable deployment, – Compatibility with many LV. Standard deployment system – Tubular frame, – Spring assisted ejection, – Payload of 3 single CubeSats (or combination with larger CubeSat-based picosats). © CalPoly © CalPoly  References (1) General references – Sutton, G. P., Biblarz, O., Rocket Propulsion Elements (8th ed), John Wiley and Sons (2010) – Brown, C., Spacecraft Propulsion, AIAA Education Series (1996) – Niederstrasse, C., Small Vehicles – A 2018 State of the Industry Survey, USU Conference on Small Satellites, SC18- IX-01 (2018) Launcher user’s handbooks – Dnepr: http://snebulos.mit.edu/projects/crm/DNEPR/Dnepr_User_Guide.pdf – Vega: http://www.arianespace.com/launch-services-vega/Vega-user's-manual.asp – Pegasus: http://www.orbital.com/NewsInfo/Publications/pegasus_ug.pdf  References (2) ISS Launches – Kawasaki, K., & , Matsumura, Y., Deploying Small Satellites from JEM Kibo, JAXA Today 6, 8–11 (2012) – JEM Payload Accommodation Handbook – volume 8: Small Satellite Deployment (https://iss.jaxa.jp/en/kiboexp/jssod) – Swartwout, M., The First One Hundred Cubesats: A Statistical View, Journal of Small Satellites 2, 213–233 (2013) Very Small Launchers – de Groote, K., Perseus : a Nanosatellite Launch System Project Focusing on Innovation and Education, in 7th International Symposium on Launcher Technologies (2007)  References (3) New Launch Methods – Clarke, A. C., Electromagnetic Launching as a Major Contribution to Spaceflight, Journal of the British Interplanetary Society 9, 261–267 (1950), reprinted in Ascent to Orbit, Wiley-Interscience (1984) – McNab, I. R., Launch to Space With an Electromagnetic Railgun, IEEE Trans. On Magnetics 39, 295–304 (2003) – McNab, I. R., Stefani, F., & Wetz, D., Access to Space using Electromagnetic Launchers, in 7th International Symposium on Launcher Technologies (2007) – Lehmann, P., Behrens, , Electromagnetic Railgun Technology for the Deployment of Small Sub-/Orbital Payloads, in 7th International Symposium on Launcher Technologies (2007) – Degtyarev, A., et al., Gun Launch System: Latest Field Testing Results, in 7th International Symposium on Launcher Technologies (2007) © Doug Drexler The lecture was way too long. Mr. Zulu, take us out of here at warp 9.