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v1.1.0 (10 June 2022) Full Screenshot Gallery Currently the mod adds 12 different engines. You can read about the engines in the Wiki. You will notice that they run on a variety of fuels. These fuels provide advantages and disadvantages compared to stock Liquid Fuel and Oxidizer, having different efficiencies, thrust ranges and storage volumes. You can check out my roadmap of engines that potentially could be added. DOWNLOADS Spacedock Github Wiki Bug Reports Dependencies (included in download, check for latest versions!): ModuleManager (4.1.4) B9PartSwitch (2.17.0) Community Resource Pack (1.4.2) DeployableEngines (1.3.0) Installation: Merge the GameData folder in the release .ZIP with the GameData folder in your KSP installation, by placing the GameData folder in to the 'Kerbal Space Program' folder. All of the dependencies come included with the mod. The "RocketMotorMenagerie" folder, as well all dependencies, should appear alongside the "Squad" folder within 'Kerbal Space Program/GameData'. To install the Extras, place any or all of the subfolders within 'Extras' into 'Kerbal Space Program/GameData'. Recommended Mods: The following mods are strongly recommended for an enhanced gameplay experience: CryoTanks (1.6.0) - This provides a set of patches provide fuel-switching features for the most basic LF/O tanks as well as orbital fuel tanks specially designed to contain cryogenic fuels Waterfall (0.6.3) - A cool way to create and drive engine effects more effectively. If you want cool engine plumes, get this mod. Frequently Asked Questions How are the parts balanced? Parts are balanced against stock KSP parts, which means the launch vehicles might seem overpowered. The parts are best suited for a 2.5x~ system rescale or JNSQ. Will you add feature/part xxx? I certainly accept pull requests. Please target all such things to the dev branch though! Or, you can make a constructive suggestion on the forum thread. Please check the roadmap above before asking. Licensing

Objectives: Establish constant presence of kerbls on Dune in long term and have regular science missions in short term. Make it as cost effective and simple as possible. Establish refinery/ refueling station on Ike for future missions aimed to explore Kerbol system and beyond. Steps: Send unmanned probes for scouting surface of Dune/ Ike for future missions. Establish constant network to KSC. sent 1st manned mission for fly-by of Dune and to test spacecraft itself. Sent 1st mission for 1st landing on Dune for exploration. Make more manned missions for scouting possible locations for establishing future bases. Repeat steps 2-5 for Ike. Establish fuel depot around Dune for future exploration missions.

I used to post on this forum in 2014, back when I was 13, and things have changed a lot since then. What never changed, though, was my love for KSP. And I decided the best way to show my love for KSP and make a return to the forums at the same time is with a mission report, as I've always loved those. A video game is about gameplay anyway, is it not? The mission in question I'm posting about is a Science Mode mission. It was a well-planned and fruitful fly-by of Minmus, that of which rewarded me with a lot of science points. Everything from ascent to landing was nominal, and I had enough delta-v thanks to using drop tanks (which I would highly recommend if you're looking for ways to get more delta-v) that, by the time I had reached a circular-ish (as in not perfectly circular) low orbit around it, I was eager to attempt to make a landing, but knew that a lack of landing legs might have spelt out stranding Jeb on Minmus, so I decided not to do that. The Experiment Storage Unit's ability to store science stuff truly came in handy on this mission, because, without it, I wouldn't have been able to store multiple experiments how you can't if you don't have it. That's how I was able to store eight experiments instead of the lower amount it would have been restricted to. The placement of the ESU indeed made the final re-entry stage taller, but it had no noticeable negative effects, as I had postulated that it wouldn't. As far as water landings go, I actually landed quite close to shore, with an impressive amount of science points to boot. Talk about a successful mission, eh? 203.1! That is no joke compared to the much more modest numbers that I used to get. In fact, the type of stuff that I can do now is what would make the KSP poster me from something like ten years ago as of writing this very impressed about it. It's moments like that which remind me why I fell in love with this game in the first place. As Jeb safely touched down near the shores of Kerbin, I couldn't help but feel a profound sense of accomplishment. The science points gathered were impressive, marking a significant leap forward in my pictured Science Mode career within KSP. Reflecting on the mission in retrospect now as I wrote this post, I can't help but marvel at the advancements in my gameplay skills since my early days of basic missions done on pirated copies of Kerbal Space Program. This time, every decision felt deliberate and informed, from the choice of using drop tanks for extended Delta-v to the careful management of scientific data with the Experiment Storage Unit. From the KSC launchpad, Jeb's spacecraft soared into the Kerbin sky, its design meticulously plotted to optimize delta-V and fuel usage, hence the drop tanks. And with the precision of a seasoned veteran Kerbonaut, Jeb soared gracefully towards Minmus in his spacecraft, fueled by the aforementioned drop tanks. The decision not to risk a landing without proper equipment showcased Jeb's commitment to safety and mission success, a decision that later proved wise as he navigated doing science work above the icy landscape of Minmus from a cautious yet cozily close orbit instead without a hitch, with more than enough delta-V and fuel to spare. The Experiment Storage Unit, smartly integrated into the capsule design, stored a treasure trove of scientific data from multiple experiments worthy of earning nods of approval from the Kerbals at the KSC back home. Especially from the scientist Kerbals over at the Research & Development Facility. Oh, and let me tell you about the return burn from Minmus! Funny thing that. Jeb's was cruising back towards Kerbin, and everything seemed on track, right? Well, the maneuver that was done for that had put me in a unique position that I was not expecting it to — a slingshot course from Kerbin to the Mun and back! Now... normally, you'd think, "Hey, free Mun fly-by, cool!" But, nope... not this time — I had over 100 points of science on board from that epic Minmus trip. And I was not about to risk losing all of that. So, instead, I slammed the brakes, metaphorically-speaking, using a retrograde burn and rerouted into a smooth re-entry trajectory instead. Sure, it would've been awesome to swing by the Mun and back, but science is science, folks. And when you've got that much valuable data packed into your capsule, you play it safe. Although Jeb may have missed out on a Mun adventure, it was ultimately way more worth having those science points safely tucked away than to risk it all. KSP's all about calculated risks, and this time, caution was definitely the name of the game. Jeb might be a daredevil in space, but I feel like even he knows when it's time to play it safe for the sake of scientific progress. Anyway... that's all. End of mission report story. If you're reading this, I want to thank you for sharing in my excitement and passion for Kerbal Space Program. My thoughts that the Kerbal Space Program community is truly something special have remained unchanged, and I'm glad that a lot of you folks are still here. I look forward being part of the KSP1 community again, as an adult with advanced knowhow now, as well as sharing more adventures with you like this all soon. I'm glad to be back, and more than curious to hear what all of you here at the forums have to say — especially after not having made a forum post before this in so long. And, plus, my fingers have started wanting to type words about the serious business that is putting cute little green creatures into space more than ever. Here's to the many more missions, discoveries, and discussions ahead. Fly safe, aim high, and may your orbits be ever stable! Cheers, Qwotty

Howdy Everybody! First off, I want to thank @Pudgemountain for this idea. He recently posted the results of an ai program attempting to recreate images of the Saturn V rocket. The results were...something to say the least. However, the images provided interesting ideas for rockets. I invite everyone on the forums to create spacecraft based on their own ai images! Post ai images for others to springboard off of, post images of your KSP creations, even both! I will also be posting ai generated images of my own so that folks can pick out ideas they want to bring to...reality?...kerbality?...whatever. Just a few rules: 1) Any craft that are created must be functional. Rockets need to be able to get off pads, spacecraft can achieve proper orbital spaceflights, etc. 2) If you decide to post your own ai inspired craft, show the original image that provided your inspiration. 3) Whatever ai generation software you use, be sure to provide credit to that site. On that note, make sure they allow public usage of images you create (shouldn't be an issue, but you never know). Above all else, have fun! Let your imagination run wild! If you want to provide a mission report or a short story for your vehicle, by all means! Below I have the original post plus examples I worked up today. Above is one of the original images that gave rise to this idea. Again, send your praise @Pudgemountain's way! Here I recreated the last rocket on the right. Then I decided to check out these ai image generators myself. Which gave me more ideas. Also I started typing a lot of random and "what-if" ideas to toy with. One of my KSP saves, 101 Rocketry, is a US-inspired private aerospace company with a bit of a Taiwanese/ROC flair. I recreated the top center rocket (at least my interpretation). Almost forgot, I have the craft files for these two rockets if you want to mess around with them: https://kerbalx.com/ManateeAerospace/XN1-Tallcom-A https://kerbalx.com/ManateeAerospace/Salamander-II

No matter what I try, my spacecraft just won't lift off. When I press space, I get what sounds like a decoupler sound effect in slow motion and then my rocket slowly falls. I think it might have something to do with the lag, since I had 3 FPS the whole time?

I require a Duna lander since I am not very good at making rockets and I can't find a good Duna lander and rocket, and it would help very much if someone could refer me to one. Disclaimer: STOCK ONLY

A pair of shuttles I built, to ferry probes to just about anywhere (Hence the name, Katurn V Probe Delivery System). Use RealFuels, or the poor thing wont have enough power to even get off the pad. With probe (uses RealAntenna) https://kerbalx.com/Natelolzzz/KVPDS Without probe https://kerbalx.com/Natelolzzz/KVPDS-No-Probe They handle like a brick, so be cautious!

Type your favorite engines out of the listed: Nerv Atomic Engine Mammoth Liquid Fuel Engine Rapier Jet Engine Mainsail Engine Vector Engine Clynsdale Solid Fuel Also (hopefully without breaking any rules) please like my thingamajig to improve my credit. ~Cheers

With Iranian rocketry progressing quickly (they're planning to launch a human into orbit by 2030), I was inspired to create this mod. This is my second mod, so feedback is appreciated! Safir: Salman: Download on SpaceDock: https://spacedock.info/mod/3550/Kermanshah Rocketry Dependencies: Waterfall License: CC-BY-NC-SA (Configs) and All rights reserved (Textures and Models)

Have you ever thought of launching rockets the wrong way? Retrograde Launch Systems gets you covered! They provide several rockets for you to launch westwards. This is my first mod, don't expect great textures or models for now. Download (SpaceDock): https://spacedock.info/mod/3545/Retrograde Launch Systems License: CC-BY-NC-SA (CFG), All Rights Reserved (Textures and Models)

In this game your objective is to improve the rocket prior to you! (Heavy inspiration from @AlamoVampire) To improve a rocket and you can use anyone of the methods below! Methods Method 1. You can post a rocket that is its next evolution. (Atlas 2 gets followed by Atlas 3A and so on) Method 2. If it uses technology used by another rocket (RS-68 for Ares V followed by Delta 4) Method 3. It looks similar ( H1 Followed by Delta II) To make sure the fun and games can continue smoothly we have to follow the rules below. Rules Rule 1. Fictional rockets are allowed to an extent (Things like FAM is allowed but things like Interstellar is not) Rule 2. You must wait 10 minutes since the last rocket you posted Rule 3. Follow the forums rules. Rule 4. Only one rocket per post. Rule 5. No spacecraft this is about rockets! Now I guess I will start us off with the first rocket to reach space the V2!

Hey! I've always been in love with atmospheric flight, but it was not until I played KSP that I found a fondness for spaceflight. I've always had a love/hate relationship with maths, i.e, I love the practical science/engineering/business applications of it, but it costs me horrors to do anything beyond basic equations. Anyway, a few weeks ago, as I was browsing the Internet, I came across a small PDF booklet that piqued my curiosity. It was titled "HOW to DESIGN, BUILD and TEST SMALL LIQUID-FUEL ROCKET ENGINES." I gave it a quick reading, skipping over most of the maths, and realized that the apparent complexity in the design of a rocket engine stems not from the engine itself, which is a relatively simple machine, but from the fact that a flight engine has to fit a very harsh set of criteria: It needs extreme levels of both thrust and efficiency. It has to be extremely lightweight, and the tanks and piping have to be lightweight too. Cost is usually not an issue, or is pretty low in the priority list. That set of criteria produces the awesome beasts we know and love, but in the process also makes them extremely complex and costly machines. (Think turbopumps, regenerative cooling, exotic materials and building techniques, cutting-edge avionics and software, ultra-precise machining, etc) I realized, that, were one to have a different set of priorities, one could take the design of rocket engines out of the realm of the true rocket engine engineers (usually teams of specialists in aerodynamics, chemistry, thermodynamics, stress analysis, avionics, and the list goes on and on) and into the hands of a single hobbyist with barely high-school math skills like me. Enthusiastic, I gave the book a more thorough reading, and found out that it was more of a "How To" guide (Insert X value into Equation 4, take it from table B, and so on), and less of a true rocket engine design book. Given the fact that I actually want to learn design instead of just blindly following along a guide, I decided upon complementing it with other bibliography, mainly "Rocket Propulsion Elements", a monster of a book at 700 pages, and filled to the brim with complex math, which, nevertheless, has managed to solve (with considerable effort and headache on my part ) all the doubts such as Why is X done in Y way?, where does this precomputed value we're told to use come from?, etc. left in the wake of the smaller book. Hence I started the design process, and am currently in the phase of producing CAD drawings for manufacturing and assembly (i.e, I'm almost done) I've decided to share the process with you in order to: Give back to this awesome community at least a tiny bit of which it has given me over the years of playing KSP. Fully review the design process from start to finish as I write this, in search of errors. Learn even more as I search for ways to explain complex concepts in forms that are simpler to grasp than mere maths. Without further ado, let's dive in! I started the design process by listing a set of criteria for the engine to meet, in order for it to be a realistic, doable project for myself. Things I want or need: Simple. Safe (Well, as safe as a controlled explosion can be anyway) Cheap to build and operate Things I do not want or need: Extreme high performance. Or any performance at all. As long as it makes a supersonic flame and lots of noise, I'm happy. Lightweight. Expensive/Hard to find/Toxic propellants. Regenerative cooling (Arguably the hardest part in the design of any rocket engine) Expensive/exotic materials. Complex/extremely precise machining of parts. Gimbaling Given that different design criteria, the project becomes a lot simpler indeed! After outlining my requirements, I made the three most basic decisions that will drive the rest of the design process. Propellants to be used. How will the propellants be fed to the engine Thrust level to be achieved. After careful consideration, and a dive into Elements of Rocket Propulsion, and some Wikipedia to check chemical properties, I settled upon Gaseous Oxygen and Methyl Alcohol as propellants. The oxidizer, gaseous oxygen (GOX) is cheap, easy to find, non toxic, non cryogenic (does not require cryogenic valves, piping, engine pre-chilling, etc), has a slightly higher performance than liquid oxygen, and it also comes pre-pressurized (No pump required). It has a big drawback, in that the required tanks and pressure regulation devices are large and heavy (think high pressure storage of a gas which uses up a large volume), and, while that would be an instantaneous No-No for an engine to be used in a flight rocket, it was unimportant for my intended use. The fuel, methyl alcohol, also known as methylene or wood spirits, was chosen because, while it is more expensive than gasoline or kerosene, it burns at lower pressures and temperatures than those, therefore making the unspoken requirement "The engine should not melt/explode" a bit easier to comply with, and it can be bought at any hardware store. Methyl alcohol is toxic, but only upon ingestion and it's not horribly toxic or carcinogenic like other propellants or oxidizers such as hydrazine, aniline, red fuming nitric acid, dinitrogen tetroxide, etc) I also decided to use a pressure fed design, as, in keeping with the simplicity premise, I want to avoid turbopumps, gas generators, and all that sort of things that make complexity, cost, and the number of potential failure points to increase. The thrust level I decided upon was 100 newtons (10.2 Kgf or around 22 lbf). It's pretty darn puny for a rocket engine, but it's a nice round number, and should still be an interesting challenge, which should be achievable without: Needing huge chamber pressures/temperatures. Having a large fuel consumption. Rocket engines are inefficient machines by nature, and I don't want to go broke after the first few minutes of operation, With that decided, it is time to determine the basic operating parameters of the engine, such as mass flow, chamber pressure, etc. that will then be used to determine the materials and physical dimensions of the engine. That is already done, but I have to review it and convert from my scribbled design notes to a good quality post. Until then, I leave you with this render of the combustion chamber / nozzle assembly as a teaser of things to come. Dec 04 2015: In the last installment, I decided upon the propellant combination and thrust target. Today, I will determine the most basic operating parameters of the engine, and, upon those, calculate some other parameters which must be known in order to start calculating the basic physical dimensions of the thrust chamber and nozzle. Also, I keep all the parameters that I determine/calculate, in a large table, that is kept handy and lets me have all the data that I might possibly need, ready at a quick glance. This is the table so far, and from now on all results will be added to it, and data for any calculations, sourced from it too. ENGINE MASTER DATA TABLE Parameter or Dimension Value Metric Imperial Propellants GOX/Methanol Thrust 10.2kgf / 100N 22.5 lbf Now, in order to get started with the physical design, we have to know a few parameters: Chamber pressure (Combustion pressure) Combustion Temperature. Mixture Ratio (Proportion of oxidizer to fuel) Approximate ISP (This is mostly a rough number dependent upon the propellant combination, and will be later adjusted to account for engine geometry losses) These parameters can be calculated, but designing an engine from scratch, with no reference numbers, is a daunting task. Fortunately there are huge amounts of precalculated data on the subject, made available by either government or private organizations, and we can easily source them from tables. As indicated on the table, these parameters are determined for expansion to 14.7 PSI, which is sea level atmospheric pressure. That is good enough for me, because this engine will not be used at very high altitude or a vacuum. (I live at 2700 ft above sea level) Also, not indicated on the table (one of the things the book assumes you to know/realize) is the fact that these pressures, temperatures, and ISP's, are based upon a stoichiometric mixture ratio (there is just enough oxidizer to burn all the fuel). Any other ratio will result in lower pressures and temperatures, and less performance, which makes sense, because you are either low on oxidizer, having unburned fuel go through the engine, and then burn with the outside atmospheric oxygen without producing useful thrust, or you have an excess of oxidizer going through the engine, and, given that there is not an unlimited amount of space inside the engine, any excess in oxidizer means a corresponding lack of fuel. (There is an exception to that if running fuel-rich reduces the molecular weight of your exhaust, such as in hydrogen/oxygen engines, but that is honestly beyond the scope of this discussion) Now we are ready to determine a few other rough parameters, most importantly, the engine mass flow rate. The engine mass flow rate will let us know the mass of propellants required for operating the engine at the desired thrust level. The formula for engine mass flow rate is: To understand why that does even make sense (It took me awhile to realize why it did, and I was very confused before that) you have to take two things into account: Mass conservation principle. No matter what is chemically happening inside the engine as propellants are burned, the same amount of mass that enters, will leave. Unless you somehow create a nuclear reaction, in which case some mass will be converted to energy, but that is a very, very, very unlikely outcome Specific Impulse (ISP) is just a fancy way of saying "Hey, this engine could develop X amount of thrust if it burned 1 unit of mass per unit of time" Thus, given that we know our thrust target and also know our rough ISP, we can proceed to calculate the amount of mass entering and exiting the engine per second. I'm pretty sure this engine will consume less than 0.1 kg of propellant per second, but let's find out the exact value. 100 Newtons are 10.1971621 kgf. Therefore our engine has a thrust of 10197 grams. Ah, metric system, how can I not love you 10197 / 248 = 41,116935483870967741935483870968 grams/sec. So, when the engine consumes 41.12 grams of propellant per second, it will emit 41.12 grams of exhaust gasses per second, and produce the 100 newtons of thrust. (In theory). Now based on that total, we will determine which part of the propellant mass is fuel and which part is oxidizer. (This will be used later in the design of the injectors, and fuel system and is better determined now than then) Given the oxidizer/fuel ratio of 1.2, as per table 1, we then can determine the mass flow ratios to be as follows: Oxidizer flow = Emfr * R / (R+1) Oxidizer flow =0.403*1.2/(1.2+ 1) Oxidizer flow = 0,2198181818181818 newtons / sec oxidizer Fuel Flow = Emfr * R /(R+1) = Fuel Flow = 0.403/(1.2+1) Fuel Flow = 0,1831818181818182 newtons/sec fuel I know I'm not supposed to use newtons as a mass unit, and later realized that mistake, but the results are the same whether I use pounds or grams , only expressed in the pertinent unit. I have no clue why this is so, and if anyone could explain, I'd be grateful. Now with these parameters calculated, we can dive into the meat and potatoes of the project, and start calculating the physical dimensions of the engine, and also you'll get to see me suffer through some more complicated maths, but that will have to wait for the next installment. Until then, this is the engine data table, with all the data that we have determined or calculated so far. ENGINE MASTER DATA TABLE Parameter Value Metric Imperial Propellants GOX/Methanol Thrust 10.2 kgf 22.5 lbf Chamber Pressure 2.068 Mpa 300 psi Mix ratio 1.2 ISP 248 s Total Mass Flow 41.1 gr/s 0.0906 lb/s Mass Flow (Oxidizer) 22.42 gr/s 0.049428 lb/s Mass Flow (Fuel) 18.68 gr/s 0.041182 lb/s If you have any insight, questions, or even better, have found an error, please let me know Dec 12 2015: Hey! After determining operating parameters, today we are going to determine some gas values, that we will then use to determine chamber dimensions, nozzle outlet diameter, expansion ratios, throat diameter, etc. Let's get started! The idea behind a DeLaval nozzle (That's how a rocket nozzle is called) is to transform a high pressure, high temperature, low velocity gas, like the combustion products, into a low pressure, relatively low temperature, and crazy-high speed gas. (Remember that momentum = mass times velocity, and given that gasses tend to be very light, in order to produce useful thrust, velocity has to be extremely high) Velocities of 2 km/s are not unheard of for small hobby engines This image shows the profile of the gases in a DeLaval Nozzle: Notice that the gasses after the throat are supersonic, and that is done in order to prevent pressure perturbations from travelling upstream (any pressure perturbations travel at the speed of sound) This is critical, because otherwise the nozzle would behave as a Venturi tube, and produce an exhaust of similar pressure and velocity as given in the inlet, which would be useless for us. Now, you'd think that calculating a diameter that will produce a desired Mach speed would be easy, but it turns out that the local sound speed (Mach number) of any gas is affected by pressure, temperature, and density... And guess what, a nozzle varies pressure and temperature along its whole length! Now the math starts to pick up in complexity! First, we have to determine the temperature of the gas in the nozzle throat (Tthroat). That is because, as explained above, the gas temperature at the nozzle throat is less than in the combustion chamber due to loss of some thermal energy during the acceleration of the gas to local speed of sound (Mach number = 1) at the throat. Gamma (the Y shaped Greek letter) is the ratio of gas specific heats, a dimensionless value (much like the Mach number), which relates to the heat capacity at a given volume for a gas. For the products of hydrocarbons and gaseous oxygen combustion, Gamma equals 1.2 Tgas = 1 / (1 + ((1.2-1)/2)) Tgas = 0.90909090909 of the Chamber temp. Tgas = 0.90909 * 3155 º K Tgas = 2868.18 º K or 2595 ºc The chamber (combustion) temperature is determined for this propellant combination from Table 1. Now, we have to determine the gas pressure at the nozzle throat.The pressure at the nozzle throat is less than in the combustion chamber due to acceleration of the gas to the local speed of sound (Mach number =1) at the throat, as given by So, Pgas = 300 psi * (1+((1.2-1)/2)) ^-(1.2 / (1.2 – 1)) I just rage-quitted there, and cheated with Wolfram Alpha. (which, by the way, is a wonderful free online tool that I recommend you check out) So, pressure at the throat is 169.34 psi. Quite a large drop, about half of the initial value, which is to be expected in this kind of nozzle. So far, so good. Also, the gas will have to be expanded to atmospheric pressure before exiting the engine. This is important for future calculations. Given that I live at 2700 ft above sea level, I just asked Wolfram Alpha which was the atmospheric pressure at that altitude. Unfortunately, time is in short availability for me right now. So, I give you the current status of our calculations. ENGINE MASTER DATA TABLE Parameter Value Metric Imperial Propellants GOX/Methanol Thrust 10.2kg 22.5 lbf Chamber Pressure 2.068 Mpa 300 psi Maximum Reaction Temperature 3155ºK 5679 ºR Mix ratio 1.2 ISP 248 s Expansion pressure 918 Mb 13.31 psi Total Mass Flow 41.1 gr/s 0.0906 lb/s Mass Flow (Oxidizer) 22.42 gr/s 0.049428 lb/s Mass Flow (Fuel) 18.68 gr/s 0.041182 lb/s Gamma 1.2 Throat Gas Temperature 2868.18ºK 5679 ºR Throat Gas Pressure 1.168 Mpa 169.34 psi Please join me in the next installment, when we determine Mach numbers and finally some physical dimensions! Until then, if you find any errors or have comments/suggestions, please do let me know. Thanks. Dec 18 2015: Hey! Real life has been hell these days! Fortunately, now I've had time to review another part of the design. Onward! Now that the gas parameters, such as temperature and pressure at the throat have been determined, and we know the mass flow of the engine, we can proceed to calculate throat area, and from that, derive throat diameter (The first physical dimension) Throat area is given by: Where R is the universal gas constant, M is the molecular weight of the exhaust gasses, and Gc is the universal gravitation constant. Athroat = ((Mflow/Pthroat) * ((R * Tthroat ) / (Gamma * gravitational constant)) ^1/2 Athroat = (0.0906 lb/sec / 169.34 psi) * ((64.388 foot-pound/pound/degree Rankine * 5679 degrees Rankine)/ (1.2 * 32.2 foot/sec^2 )) ^1/2 Athroat = (0.0906 lb/sec / 169.34 psi) * ((64.388 foot-pound/pound/degree Rankine * 5679 degrees Rankine )/ (1.2 * 32.2 foot/sec ^2 )) ^1/2 Athroat = (0.0906/169.34 psi) * ((64.388 *5679)/ (1.2*32.2)) ^1/2 Athroat = (0.0906/169.34 psi) * (365659.452 / 38.64) ^1/2 Athroat = (0.0906/169.34) * (365659.452 / 38.64) ^1/2 Athroat = (0.0906/169.34) * (365659.452 / 38.64) ^1/2 Athroat = 0.0520461 square inches, or 33.5781 square mm Given this area, we can proceed to determine diameter, by simple geometry of circles. Dthroat= 4*33.5781 / 3.14159265 Dthroat= 4*33.5781 / 3.14159265 Dthroat= 6.53858 mm I'm sorry guys, I really wanted to push out more content today but an unexpected work issue has arisen (yet again) *Sigh*.. I'll have a fuller update ASAP. Sorry for the really crappy little update, but hey, progress is progress! PS: Almost forgot, this is our Data Table now: ENGINE MASTER DATA TABLE Parameter Value Metric Imperial Propellants GOX/Methanol Thrust 10.2kg 22.5 lbf Chamber Pressure 2.068 Mpa 300 psi Maximum Reaction Temperature 3155ºK 5679 ºR Mix ratio 1.2 ISP 248 s Expansion pressure 918 Mb 13.31 psi Total Mass Flow 41.1 gr/s 0.0906 lb/s Mass Flow (Oxidizer) 22.42 gr/s 0.049428 lb/s Mass Flow (Fuel) 18.68 gr/s 0.041182 lb/s Gamma 1.2 Throat Gas Temperature 2868.18ºK 5679 ºR Throat Gas Pressure 1.168 Mpa 169.34 psi Throat Area 33.5781 mm2 0.0520461“2 Throat Diameter 6.53858 mm 0,2574244 “ Feb 18 2015: Hello guys! Sorry I left all of you hanging in there, but I've been having all kinds of Real Life Stuff™ going on! I can't promise updates will be regular anymore, but this project is in no way shelved or anything. In the last installment, we had determined the gas pressure, temperature, and throat area of the nozzle. Now, with that data on hand, we can proceed to calculate the best bell end diameter that will provide expansion to the desired pressure and prevent the engine from running under or overexpanded. (Don't worry, I'll explain these terms in a second) In order for us to understand why expanding to a predetermined pressure is important, you have to go back to the definition of a DeLaval nozzle that I posted some paragraphs above. "The idea behind a DeLaval nozzle [...] is to transform a high pressure, high temperature, low velocity gas, like the combustion products, into a low pressure, relatively low temperature, and crazy-high speed gas." So the nozzle does useful work (accelerating a gas) by taking energy from its heat while reducing its pressure. I never even thought this would be significant, I mean, the larger the expansion, the better the performance you extract from the engine, right? But as a thought experiment, I decided to imagine the "perfect" engine. This perfect engine would have an infinitely large exhaust nozzle, it would drop the exhaust pressure to zero and the exhaust temperature to absolute zero, and thereby convert all the available heat from the exhaust gasses into kinetic energy. Exhaust velocity would NOT be infinite, because there's only a limited amount of heat energy to begin with, and, given the infinitely large nozzle would also be infinitely heavy, that would render our perfect engine useless, but hey, this is only a thought experiment in a perfect vacuum... And then it hit me, that in fact a real engine would not operate in a perfect vacuum, where the ideal exhaust pressure is zero, but it would operate inside an atmosphere, where the ideal expansion is to atmospheric pressure. To better understand why this is so: Imagine you are sitting with your engine at sea level. Therefore, the pressure of the engine exterior is 1 atmosphere, or 14.7 psi. Now imagine you had 300 psi in the combustion chamber, and your hypothetical nozzle had been designed to reduce pressure at the exit to 100 psi. So, what happens when you start said engine? Your nozzle works as expected, and it reduces exhaust pressure to 100 psi, with a proportional temperature drop. Then, once the gasses leave the nozzle, what happens? They immediately proceed to expand to 14.7 psi, further cooling in the process. Therefore your nozzle is underexpanded, and it is wasting gas energy (Remember, any gas that expands outside the engine is useless for thrust, much like excess fuel would be (There is an exception to that if running fuel-rich reduces the molecular weight of your exhaust, such as in hydrogen/oxygen burning engines, but that is honestly beyond the scope of this discussion)). Now to the opposite end of the spectrum: Imagine you take the same engine and change the nozzle for one that goes to, say, 0.5 psi. As the gasses go further down the nozzle, their pressure will decrease, until it matches that of the atmosphere. At said point, they stop expanding, because the atmospheric pressure exerts a force equal and opposite to that of the inner gas pressure, and the exhaust will form a column that is "pinched" by the atmosphere and will exit the bell without expanding any further. This seems like it would be good enough, right? You get a slightly larger and heavier nozzle, but for that price, you make absolutely sure that you're expanding the gas as much as it can expand, and getting all the thermal and pressure energy you can get out of it. The exhaust is as cool as it can get, it's at ambient pressure, and you've extracted all the velocity you can extract. Then who cares if the nozzle is a bit too large? Well, in an ideal world that would be OK, but in the real world, having parts of the nozzle not filled with exhaust is a bad, bad idea. The best that can happen is that the gas "sticks" to the nozzle walls after its expansion is done, you get vacuum "bubbles", Mach diamonds, turbulence, etc. in the exhaust and you lose thrust. (That happens with mild overexpansions) and the worst that can happen is the flame flopping around like crazy and banging the nozzle walls randomly until the vibration, noise, and mechanical stress of the turbulent gas flow cause the engine to experience R.U.D. (Rapid Unscheduled Disassembly) Real rockets have a problem with that. Especially first stages! First stages have to go from sea level to almost a vacuum! So how do they avoid gross underexpansion or overexpansion? Well, by compromising, and using a nozzle that is designed to work halfway between sea level and vacuum. So upon start up they are overexpanded, and as they climb they reach their design altitude (perfectly expanded), and then past that they become underexpanded. Example overexpanded nozzle. You can see the telltale Mach disks. And my favorite underexpanded one, Saturn V going uphill You can check out the expansion of exhaust gasses in this video of the Mars Climate Orbiter launch. Check out how big that plume gets as the atmosphere gets thinner and thinner. That was when I came across what I thought was the simplest engine design calculation so far: With said constant already being helpfully provided by the author. But alas, I'm always curious, and I dived into Rocket Propulsion Elements, to find out why relative gas expansion was so simple. Oh, how I was to repent. Turns out, said constant is only valid for sea level. For expansion to a different pressure, you need either a new constant, or you need to do math of the kind that gives you chills. Nevertheless, once I was in, i had no choice but to press forward (Just kidding, I had fun learning about it) These equations will be used to calculate the Mach speed of the exhaust gasses, and once we have that, find an exhaust area that will yield exhaust pressure equal to the local atmospheric pressure for that Mach number. Once again, Wolfram Alpha proves to be an invaluable tool for the hobbyst rocket engineer who wants to save time and headache. An exhaust velocity of Mach 2.62 sounds incredibly high, but actually, is pretty much on the lowest end of what you will get with a rocket engine. Now that we know the area of the nozzle end, we can use simple circle geometry to calculate a diameter (It's the same formula we already used to derive nozzle throat diameter from nozzle throat area) Dexhaust=0.515609 inches or 13.0965 mm. Therefore our Master Data table now looks like this: ENGINE MASTER DATA TABLE Parameter Value Metric Imperial Propellants GOX/Methanol Thrust 10.2kg 22.5 lbf Chamber Pressure 2.068 Mpa 300 psi Maximum Reaction Temperature 3155ºK 5679 ºR Mix ratio 1.2 ISP 248 s Expansion pressure 918 Mb 13.31 psi Total Mass Flow 41.1 gr/s 0.0906 lb/s Mass Flow (Oxidizer) 22.42 gr/s 0.049428 lb/s Mass Flow (Fuel) 18.68 gr/s 0.041182 lb/s Gamma 1.2 Throat Gas Temperature 2868.18ºK 5679 ºR Throat Gas Pressure 1.168 Mpa 169.34 psi Throat Area 33.5781 mm2 0.0520461 “2 Throat Diameter 6.53858 mm 0,2574244 “ Exhaust gas velocity (Mach) 2.62167 Nozzle exit area 134.71 mm2 0.2088 “2 Nozzle exit diameter 13.0965 mm 0.515609 “ Join me in the next installment, where we'll calculate the combustion chamber parameters, and we will be then ready to begin sketching the innards of the chamber + nozzle. Until then, thanks for your time & patience in dealing with my ramblings, and as always, if you find a mistake, please DO let me know. I happen to dislike explosions if I have to pay for the exploding stuff. Mar 3 2015: Hi! Finally found a bit of free time! Real Life keeps me busy, and usually at the end of the day I'm too knackered to do anything other than crawl into bed .... But enough of my whining! You're here for the possible explosions rocket engine design theory. Given that we now know the throat diameter, and exit diameter, one would think that it's already time to calculate nozzle inlet diameter, but, a quick bit of thinking reveals that the nozzle inlet and chamber outlet are one and the same, so we'll kill two birds with a single stone, and calculate chamber dimensions which we can then use to derive nozzle inlet diameter. We will start by calculating the volume of the chamber, and, knowing that volume, we can make an educated guess about length/diameter ratio, and calculate exact values from there. What would a good volume be? A good volume would be one that ensures adequate mixing, evaporation, and complete combustion of propellants by the time they reach the nozzle inlet. That is so, because the nozzle is designed to work with a specific inlet pressure and temperature. Any propellant that goes past the nozzle inlet, will probably burn in the nozzle, which is a bad idea because temperatures at the throat are already pretty critical (despite being at lower temperatures, the throat is the area with less dissipation surface available, and therefore more susceptible to heat damage) and also it would throw off our pressure and temperature ratios for all the points along the nozzle, and if the chamber is too big, the gasses will have time to cool before they enter the inlet, thus reducing engine performance. So, in resume: Chamber too big: Colder inlet temperatures, performance wasted. Heavier engine. Somewhat easier cooling due to lowered gas temps at the nozzle. Risk of combustion instability. Chamber too small: Dangerously hotter nozzle, performance wasted. Lighter engine. Calculating the aerothermochemodynamics of complex hydrocarbons reacting while changing their state, pressure, mixture ratios, temperature, movement speed, and several other variables, in order to ensure complete combustion, is an awful, hellish nightmare. Trust me, I have looked at it. But turns out, there's a cheat for that. Even Real Life Rocket Scientists™ happen to use it for preliminary designs. It's called "characteristic chamber length" and is defined as the length that a chamber of the same volume should have if it were a straight tube and had no converging nozzle section. Characteristic chamber length, L* or L star, is determined experimentally for different propellant combinations, throat diameter, and combustion pressures, and it can be sourced from tables. For an hydrocarbon burning engine like mine, L* is between 50 to 70 inches. The variation is to account for injector design (propellant mixing) I decided to go with 60 inches. Vchamber = 60 * 0.0520461 cubic inches, therefore Vchamber = 3.122766 in3 or 51.173 cm3 To derive chamber length from volume, we also have to know a diameter. A good diameter for combustion chambers is around 5 times throat diameter. Dc = 5 Dthroat Dc = 5 * 6.53858 Dc = 32.6929 mm – 1.287122 inches This is for the cylindrical portion of the chamber. For a small chamber, we can just assume the convergent segment to be 1/10th of the chamber volume, and be done with it. For the chamber area, i just went with my trusty ally, Wolfram Alpha. Lc = Vc / (1.1 * Ac) Lc = 3.122766 in3 / (1.1 * 1.3012 in2) Lc = 3.122766 / (1.1 * 1.3012) Lc = 3.122766 / 1.43132 = 2.18174 inches - 55,416196 mm And thus, our Engine Master Data Table is beginning to fill with physical dimensions. ENGINE MASTER DATA TABLE Parameter Value Metric Imperial Propellants GOX/Methanol Thrust 10.2kg 22.5 lbf Chamber Pressure 2.068 Mpa 300 psi Maximum Reaction Temperature 3155ºK 5679 ºR Mix ratio 1.2 ISP 248 s Expansion pressure 918 Mb 13.31 psi Total Mass Flow 41.1 gr/s 0.0906 lb/s Mass Flow (Oxidizer) 22.42 gr/s 0.049428 lb/s Mass Flow (Fuel) 18.68 gr/s 0.041182 lb/s Gamma 1.2 Throat Gas Temperature 2868.18ºK 5679 ºR Throat Gas Pressure 1.168 Mpa 169.34 psi Throat Area 33.5781 mm2 0.0520461 “2 Throat Diameter 6.53858 mm 0,2574244 “ Exhaust gas velocity (Mach) 2.62167 Nozzle exit area 134.71 mm2 0.2088 “2 Nozzle exit diameter 13.0965 mm 0.515609 “ Chamber Volume 51.173 Cm3 3.122766 “3 Chamber Diameter 32.6929 mm 1.287122 ” Chamber Area 839.5 mm2 1.3012 “2 Chamber Length (including Convergent Segment) 55,416196 mm 2.18174” Please join me next time, were we'll calculate chamber walls, dabble in safety margins, and make a first crude sketch of the engine (Spoiler: It does end up looking like a rocket engine) Until then, if you happen to find any errors, or have feedback, please do so. Thanks Apr 30 2016: Wow! It's been a long time! Sorry for the delay guys... Real life has been absolutely hectic, work issues, study issues, family issues, you name it you got it! Despite the long time between updates this project is not dead at all and I've been itching to show some of the progress I've made. So, without further ado, let's dive in! In the last installment, we had finished determining chamber and nozzle dimensions, but these are the inside ones only, and now we will calculate wall thickness. Every point in the chamber and nozzle has to be strong enough to resist the pressures involved, otherwise the engine will explode. I've decided that, in order to simplify the design, I will simply use a constant wall thickness, suited for the highest pressure area. This is really overkill for parts of the nozzle where the pressure will be lower, and makes the engine significantly heavier, but greatly simplifies both design and machining. Thus I shall design a vessel that can contain 300 psi with an adequate safety margin. Given that the nozzle will be automatically overbuilt, due to its lower operating pressure, I will treat the chamber as a pipe and thus greatly simplify calculation. The equation for the stress on the wall of a tube is: where S is the stress on the pipe wall, P is Pressure, D is Diameter and Tw is the wall thickness. Thus, if we replace S with the ultimate strength of our material, we can calculate the minimum wall thickness. I choose copper, given that it has excellent thermal conductivity, is easy to machine, and is cheap. The ultimate strength of copper is around 10.000 psi, but I will use a conservative 8000 psi in this calculation. S= P * D / 2Tw Tw = P * D / 2S Tw = 300 psi * 1.287122 inch / 16000 Tw = 300 * 1.287122 /16000 Tw = 0.0241335375 inch or 0.61299185 mm Of course this is the absolute minimum value, and while going with 2 mm wall thickness should be more than enough, there are other things to consider, machinability being a top priority since I don't want this project to be unnecessarily hard to machine (Machining a nozzle with walls of that thickness, in copper, will be very hard to do without deforming it) Therefore, I will make an educated guess and use a 5 mm wall thickness, which should be easy to obtain. That also gives me an 815% safety margin. This baby may melt, but an explosion is now an extremely unlikely outcome. (Thankfully) Obviously this just made the engine a lot heavier, but, then again, I don't care about weight. Now that we know all our dimensions, we need to determine our half angles, or the angles of the lines that join inlet, throat, and outlet, thus conforming the nozzle walls. For this small engine, adding a bell shape would give me major machining headaches, and produce only a minor performance improvement. Based on a simpler geometry proposal by @A Fuzzy Velociraptor, I decided to go with 15º and 40º half angles, jointed by rounded unions. I proceeded to fire up my favorite CAD software and did a quick sketch. (All dimensions in mm) I don't know about you, but to me, that definitely looks like a rocket engine. What do you guys think? Next up: We will tackle the issue of cooling. Hopefully tomorrow. No promises. ENGINE MASTER DATA TABLE Parameter Value Metric Imperial Propellants GOX/Metanol Thrust 10.2kg 22.5 lbf Chamber Pressure 2.068 Mpa 300 psi Maximum Reaction Temperature 3155ºK 5679 ºR Mix ratio 1.2 ISP 248 s Expansion pressure 918 Mb 13.31 psi Total Mass Flow 41.1 gr/s 0.0906 lb/s Mass Flow (Oxidizer) 22.42 gr/s 0.049428 lb/s Mass Flow (Fuel) 18.68 gr/s 0.041182 lb/s Gamma 1.2 Throat Gas Temperature 2868.18ºK 5679 ºR Throat Gas Pressure 1.168 Mpa 169.34 psi Throat Area 33.5781 mm2 0.0520461 “2 Throat Diameter 6.53858 mm 0,2574244 “ Exhaust gas velocity (Mach) 2.62167 Nozzle exit area 134.71 mm2 0.2088 “2 Nozzle exit diameter 13.0965 mm 0.515609 “ Chamber Volume 51.173 Cm3 3.122766 “3 Chamber Diameter 32.6929 mm 1.287122 ” Chamber Area 839.5 mm2 1.3012 “2 Chamber Length + Convergent segment 55,416196 mm 2.18174” Chamber Wall thickness 5 mm 0,19685” Nozzle Half-Angle 15º Nozzle inlet Half-angle 40º May 06 2016: Did I say tomorrow? I totally meant in a week or so! Let's get started on cooling, shall we? In order to understand the cooling needs, we first have to understand how the heat flows through a rocket engine. Most of the heat of combustion is either used up accelerating the gasses, or leaves with the exhaust, while a part of it is transferred to the chamber wall, propellant injectors, and nozzle. Heating is a problem because it can debilitate the metals of the chamber to the point at which they cannot resist the chamber pressure anymore, causing deformations which are usually followed by RUD. Therefore, we can devise of several methods to keep the temperatures within reason. No cooling at all: Use the thermal mass of the engine as a heat sink, then radiate the heat away while the engine is off. Pros: Simplest method - Cons: Run time very constrained. Passive cooling: Use either the engine nozzle or chamber walls exposed to the atmosphere as radiators. Pros: Very reliable - Cons: Complex design, a large run time requires more radiating surface than may be available, and thus, the run time is still limited without adding heavy radiator vanes. Active cooling: Use a cooling fluid circulated against the walls to get heat out of the engine. Pros: Unlimited run time. Potential to be extremely lightweight, if regenerative cooling is used (Regenerative cooling means that propellant doubles as cooling fluid) Cons: Complex design. I shall use Active cooling for this engine, for the following reasons: Safety I: If I design the engine for unlimited run time, the chance of destroying it in a 5 second initial run is extremely low. Safety II: The cooling jacket doubles as a shrapnel shield, and protects the test stand equipment from a possible explosion. It should be an interesting and educative challenge, but not a hardcore one like regenerative cooling. Active cooling works like this: (in this example, the cooling fluid is water) Small hobby rocket engines have an average heat transfer from the hot gasses to the chamber walls of about 0.5 Kw/cm2/sec, or 3Btu/sq inch./sec. Therefore, and assuming a perfect wall conductivity, this is the amount of heat that has to be removed from each square cm of the engine. Now in order to know the total heat transfer per unit time, we have to determine the inner surface area. In order to simplify calculation, I will ignore fillets and treat the engine as a cylinder for the chamber, a truncated cone for the nozzle's convergent section, and another truncated cone for the divergent section. Atotal= Achamber + Anozzle convergent + Anozzle divergent The formula for the surface area of a cylinder is: I shall modify this formula, because I do not want the total area, I only want the area of the side wall + top (the injector plate) The bottom area is shared with the convergent section of the nozzle and there is no material there to absorb heat. Therefore, Achamber = 2 * 3.14159265359 * 16.345 ^ 2 + 2 * 3.14159265359 * 16.345 * 40 So, the area of the chamber inner side walls plus injector plate inner side: Achamber = 5786 sq milimeters. The lateral area of a truncated cone, is as given by: Thus, for the convergent segment of our nozzle, Anozzle c = 3.14159265359 * (16.345 + 3.408) * Sqrt ( (16.345 - 3.408)^2 + 15.838) We use lateral area because the "bottom" of the truncated cone is the chamber radius and is not in contact with the walls, and the "top" is the throat radius, and, as such, also not in contact with walls. Therefore, Anozzle c = 840 sq mm And now, the same for Anozzle d Atotal= Achamber + Anozzle c + Anozzle d Atotal= 5786+ 840 + 145 Atotal= 5786+ 840 + 145 = 6771 square mm, or 67.71 square cm, or 10.5 square inches. The total heat transfer, "Q", is equal to the heat transfer rate "q" times the surface area of the inner walls. Therefore Q = qA Q = 0.5Kw/cm2/sec * 67.71 cm2 And thus, the total heat transfer of the engine is 33.85 Kw, or about 45 horsepower... (For the Imperial guys, about 31.5 BTU/sec) Join me next time, when we will attempt to find out exactly how much water flow does it take to get these insane amounts of heat out of the engine! If such a small engine produces these amounts of heat, my respect for the guys and gals that work on the real deal with regenerative cooling has multiplied hundredfold. May 16 2016 A few days ago, we calculated the amount of waste heat that the engine would output when working, and now we need to devise a means to get said heat out of the engine, in order to keep the operating temperatures as low as possible. Injector cooling is not an issue, as they are cooled by the inflow of propellant. Injector plate and chamber, however, are. For the sake of simplicity, I will stick to using water as coolant. Therefore, the system now has a few defined constraints: The coolant fluid must not boil. I will use water as coolant, for its high specific heat, and availability The system must be more capable than strictly needed. I don't care about mass and therefore I will have ample safety margins. Coolant flow speed of 10 m/sec or around 30 fps The coolant shall enter near the nozzle, flow all the way around the chamber, and leave near the injector plate. The amount of water mass flow (mass/sec) needed can be calculated, given the desired temperature rise and the heat input to the fluid, as given by: This is a simplified equation that only will work for water. For other cooling fluids, you need to factor in specific heat capacity. A good ΔT could be 20 ºC, that way water entering the cooling system at ambient temperature, about 20 ºC, would leave at 40 ºC, and thus a 60 ºC margin would remain before its boiling point. (68 to 108 ºF, 42.22ºF ΔT,) Wm = 31.5 / 40 Wm = 0.7875 pounds/sec, or 357 grams/second of coolant fluid. Another cool thing about using water is that, given a density (σ) of 1kg/lt, we now also know that the engine will need 0.357 liters of water per second in order to operate. (That is around 21.5 liters per minute, or 1290 liters per hour.) Now we have to calculate a pipe of such area as to obtain the desired water flow at the desired flow velocity (10 m/s should be more than enough to prevent boiling for this engine). To simplify calculation, I will treat water as a perfectly incompressible fluid. To obtain the desired mass flow at the desired velocity, the cooling jacket must have an area Ajacket, given by: The cooling jacket will therefore be like a ring around the outside of the chamber walls, with cross-sectional area Ajacket , as given by: where D2 is the inner diameter of the outer jacket and D1 is the outer diameter of the combustion chamber, given by: D1 = Dc +2Tw Where Dc is Chamber inner diameter and Tw is the wall thickness. Now we substitute and solve as this: And thus: D2 = sqrt(4mw/(Vw ^ density ^pi) + D1 ^2) D2 = sqrt(4*0.357kg /(10 m/s ^ 1 kg/lt ^3.14159265359) + 42.69 ^2 ) So, 44.33 mm is the inner diameter of the cooling jacket. I will just round it up to 46 mm for ease of machinability. That will increase coolant consumption without significantly improving cooling, but I don't care about that. Please join me in the next installment, when we finish up the coolant jacket design, including yet again safety margins, and some weird math! Until then, I leave you our ENGINE MASTER DATA TABLE Parameter Value Metric Imperial Propellants GOX/Metanol Thrust 10.2kg 22.5 lbf Chamber Pressure 2068 kpa 300 psi Maximum Reaction Temperature 3155ºK 5679 ºR Mix ratio 1.2 ISP 248 s Expansion pressure 918 Mb 13.31 psi Total Mass Flow 41.1 gr/s 0.0906 lb/s Mass Flow (Oxidizer) 22.42 gr/s 0.049428 lb/s Mass Flow (Fuel) 18.68 gr/s 0.041182 lb/s Gamma 1.2 Throat Gas Temperature 2868.18ºK 5679 ºR Throat Gas Pressure 1168 kpa 169.34 psi Throat Area 33.5781 mm2 0.0520461 “2 Throat Diameter 6.53858 mm 0,2574244 “ Exhaust gas velocity (Mach) 2.62167 Nozzle exit area 134.71 mm2 0.2088 “2 Nozzle exit diameter 13.0965 mm 0.515609 “ Chamber Volume 51.173 cm3 3.122766 “3 Chamber Diameter 32.6929 mm 1.287122 ” Chamber Area 839.5 mm2 1.3012 “2 Chamber Length + Convergent segment 55,416196 mm 2.18174 ” Chamber Wall thickness 5 mm 0,19685” Nozzle Half-Angle 15º Nozzle inlet Half-angle 40º Average wall heat transfer 0.5 kw/sec/cm2 3 Btu/sec/“2 Total inner surface area 67.71 cm2 10.5“2 Total heat transfer 33.85 kw/sec Coolant fluid Water Coolant fluid ΔT 20º C 42.22º F Coolant mass flow 357 grams/sec 0.7875 lb/sec Coolant flow volume 0.357 liters/sec 12.07 fl oz/sec Coolant density 1kg/lt 62.43 lb/ft3 Coolant flow velocity 10 m/s 32.81 ft/sec Coolant jacket inner diameter 46 mm 1.811” June 29 2016 Man, time sure flies when you're having fun horribly busy! On with the show! In the last installment, we had almost finished the cooling jacket, but some dimensions still have to be known, such as jacket inlet/outlet diameters, and jacket wall thickness. I shall use a single outlet, and two offset inlets, in order to produce a swirling motion of the coolant that should help prevent hot spots. In order to avoid pressure variations, and to keep flow speed constant, I shall keep a constant area between inlets, jacket, and outlet. The jacket has to withstand the coolant pressure, but it also doubles as shrapnel shield in case of engine RUD, and thus I will simply go for an overkill 5 mm wall thickness for the jacket, which gives us an outer diameter of 56 mm. The area of the inlets equals 1/2 of the area between the chamber outside wall and the jacket inner wall. This, as given by the area of a circle formula, equals 5221 mm2 for the jacket, and 4499 mm2 for the chamber. Thus, the coolant flow passage area is 722 mm2. and the outlet is 3.032 cm in diameter, while the inlets are half that. I'll just round it to 30 and 15 mm, for ease of machining. I'm starting to feel that the extra area I've added is counterproductive, as the design might be wasteful of water. Although better safe than sorry. I'll stick to those dimensions, and if there's excessive cooling I can simply reduce flow. And with that, the cooling design is done. Next up: Injectors! Oh boy! ENGINE MASTER DATA TABLE Parameter Value Metric Imperial Propellants GOX/Metanol Thrust 10.2kg 22.5 lbf Chamber Pressure 2068 kpa 300 psi Maximum Reaction Temperature 3155ºK 5679 ºR Mix ratio 1.2 ISP 248 s Expansion pressure 918 Mb 13.31 psi Total Mass Flow 41.1 gr/s 0.0906 lb/s Mass Flow (Oxidizer) 22.42 gr/s 0.049428 lb/s Mass Flow (Fuel) 18.68 gr/s 0.041182 lb/s Gamma 1.2 Throat Gas Temperature 2868.18ºK 5679 ºR Throat Gas Pressure 1168 kpa 169.34 psi Throat Area 33.5781 mm2 0.0520461 “2 Throat Diameter 6.53858 mm 0,2574244 “ Exhaust gas velocity (Mach) 2.62167 Nozzle exit area 134.71 mm2 0.2088 “2 Nozzle exit diameter 13.0965 mm 0.515609 “ Chamber Volume 51.173 cm3 3.122766 “3 Chamber Diameter 32.6929 mm 1.287122 ” Chamber Area 839.5 mm2 1.3012 “2 Chamber Length + Convergent segment 55,416196 mm 2.18174 ” Chamber Wall thickness 5 mm 0,19685” Nozzle Half-Angle 15º Nozzle inlet Half-angle 40º Average wall heat transfer 0.5 kw/sec/cm2 3 Btu/sec/“2 Total inner surface area 67.71 cm2 10.5“2 Total heat transfer 33.85 kw/sec Coolant fluid Water Coolant fluid ΔT 20º C 42.22º F Coolant mass flow 357 grams/sec 0.7875 lb/sec Coolant flow volume 0.357 liters/sec 12.07 fl oz/sec Coolant density 1kg/lt 62.43 lb/ft3 Coolant flow velocity 10 m/s 32.81 ft/sec Coolant jacket inner diameter 46 mm 1.811” Coolant flow passage area 722 mm2 1.119”2 Coolant inlets diameter 15 mm 0.5906” Coolant outlet diameter 30 mm 1.181” Mar 10 2017: Not abandoned! It may take me a long time, but this project will be finished come hell or high water! It's been a long time, so I'd recommend that you read from the beginning as a refresher. With that said, let's proceed. So, where was I? Ah, yes, injectors, injectors. The function of an injector is to take high pressure propellants from the feed lines, meter the appropriate amount of each (much like a carburetor), and inject them into the chamber in such a way that they can properly and efficiently burn. There are several kinds of injectors, impinging, showerhead, hollow post, pintle, etc. For this design, I shall use an impinging design. It's easy to design and build, and, while it has several disadvantages (Less efficient, very hard to throttle, small variations in shape cause big mixture irregularities, etc), these disadvantages are irrelevant to the type of engine that I'm designing. There are several "eyeballed" parameters. 100 PSI pressure drop. This should be enough to help prevent instability without requiring structural reinforcement. 20 m/s injection velocity. I was unable to find data on how an injection velocity is chosen for different propellants, however, this value is mid of the range for small hydrocarbon/oxygen engines We can now proceed to determine injector hole area, based on the physical characteristics of the propellants. Ethanol can for all practical purposes be considered incompressible. Thus, the injection area that satisfies the mass flow and injection characteristics is given by Where m is the propellant flow mass, c is the discharge coefficient, δ the density, and Δp the pressure drop. A typical discharge coefficient for round hole, small size injectors with a larger fuel manifold behind is about 0.7 The density of ethanol is about 0.75 g/cm3 at ambient pressure, and almost does not change with pressure. Pressure drop will be 100 psi. And also the bibliography I'm using (For those of you crazy cool enough to attempt a similar project) Title Author Editor DESIGN OF LIQUID PROPELLANT ROCKET ENGINES Dieter K. Huzel and David H. Huang Rocketdyne Division, North American Aviation HOW to DESIGN, BUILD and TEST SMALL LIQUID-FUEL ROCKET ENGINES Leroy J. Krzycki ROCKETLAB / CHINA LAKE, CALIFORNIA MECHANICS AND THERMODYNAMICS OF PROPULSION Philip G. Hill and Carl R. Peterson Addison-Wesley Publishing Company Ignition!: An informal history of liquid rocket propellants John D. Clark Rocket Propulsion Elements 7th Edition GEORGE P. SUTTON and OSCAR BIBLARZ JOHN WILEY & SONS, INC If you have any insight, questions, or even better, have found an error, please let me know

Go everywhere, but now with rovers! Total Mass: 2903t Dry Mass: 573t Initial TWR 1.68 DeltaV 24k Launch it from runway. If you have some problems related to the craft you can report your problem to me. Known issues: - Problem with decoupling nuclear to liquad stage. Temporal solution: time warp stages from each other. https://drive.google.com/file/d/1p8kHtaQ9tDR7ajhAe0WuQzsEAu_3fruz/view?usp=sharing If you have some problems related to the craft you can report your problem to me.

I tried to build a new rocket to go to the Mun northpole this morning. I got big problems. With the ship, it was a lousy build, but mainly with the save ships and save game commands. The saved ships that I had made were first not available ingame. (Which meant that the saves got corrupted and at one point i could not even save the game) That was the first thing that happened. When i looked in workspaces folder they were still there. The last save had the message that the rockets were unavailable. I could not save the Luna1 ship i was working on in the assembly hall. After restarting the game the ships came back and i managed to make a lift off. After going back to the vehicle assebly hall I loooked at the saved ships and they were all there. I decided to erase one of the ships and now I got back to the problem i had earlier. All ships disappered ingame and all saves had the wording that the ships were unavailable. I could not save the ship that I was working on either. I could make new saves but with the message that the vehicle was unavailable. I looked in the workspaces folder and all Luna and Mun ships had disappered. Only my crappy first ship was still there. And i could not save the ship that was in the Vehicle assembly hall. This is how it looked like. Now I am reinstalling the game from scratch. I will write here again later today if the problem remains. For my specs: https://pasteboard.co/8LMZ0j6nm6vE.png I have GTX 3070

I am working on a mod for the Neutron Rocket made by Rocket Lab. It is still WIP I am currently working on the second stage and fairings.For some reason the colliders got completely shifted and the same with the node and the fairings don't decouple. I am open to any and all ideas. I am gonna take A LOT of creative liberty since we barely know anything about this rocket. Also I need someone to help me out with the nodes and modules for the fairings.Here a few development pics: The part will be temporarily placed in the thermal section just to make them easy to find in development These are the upper stage pics:

Meet the Auora, a new advanced space craft that was made in secret development 2 months ago. Now, Andrew the Astronaut Industries is happy to present it download link: https://kerbalx.com/Andrew_the_Astronaut/Auora -Andrew the Astronaut industries PR team manager

Hey there! Im playing on ps4 and im trying to assemble my Space craft using a single high isp rocket with a lander attached Apollo-style. Thing is, everytime i tried it in the past, i remember the whole craft started wobbling in every direction and i'd like to avoid that as much as possible. Can anyone help me plz ?

i try to launch the boosters but instead it decouples and i dont know how to fix it https://imgur.com/a/iK918Q4

is proud to present.... The Zenit 3SLB launch system! This mod was originally created by the wonderful @stubbles, who at that time created one of the best rocket packs out there. This pack contains the Zenit-3SLB booster, updated for KSP 1.2.X by @MeCripp, who also included support for Animated Decouplers and Procedural Fairings. Unfortunately, after the release of v1.2, meant for KSP 0.23.5, all work on this project was terminated. After a long hiatus, @MeCrippand I are excited to bring this mod back to the KSP Community. The Zenit booster is a medium-heavy launch vehicle, capable of carrying heavy payloads to Geostationary orbit. This Zenit can bring ~10-15T to Geosynchronous orbit. It is compatible with FAR, KWRocketry, MechJeb, and DeadlyReentry. For more information about Zenit: Click here! Download from SpaceDock License: Non-Commercial CC BY-NC Future Plans: Mods included with this download: REQUIRED (If you want to use the craft file; I'll make one without the struts at a later date) Recommended mods (soft dependencies) HAPPY LAUNCHINGS!

Your objective: Build the smallest rocket possible that can get into orbit. There are 2 categories, each with 2 sub-categories. I will judge each rocket in 2 separate ways, cost and mass. Each one will have its own leader board. The rules are simple. No cheating with Hyperedit, debug menu, ect. The rocket must be completely stock. No mods that alter physics and no clipping parts to exploit physics. The rocket can be manned or unmanned, each will have their own leader board. Make sure to post pictures showing your rocket in flight and showing it's mass and cost. ALSO: The rocket must be a rocket, not a cross between a rocket and a plane. It must take off from the launch pad ,and have other rocket qualities. https://imgur.com/a/swbkZaD This is my entry LOWEST MASS ROCKETS Unmanned: 1. Pds314- 0.855 tonnes 2. Pds314- 1.091 tonnes 3. mystified- 1.491 tonnes 4. Johnster_Space_Program- 3.420 tonnes 5. DRAG0Nmon- 4.460 tonnes Manned: 1. Aetharan- 1.990 tonnes 2. Aetharan- 2.053 tonnes 3. Aetharan- 2.085 tonnes 4. Aetharan-2.159 tonnes 5. Ironbars- 2.175 tonnes LOWEST COST ROCKETS Unmanned: 1. Pds314- 640 kerbucks 2. Pds314- 2110 kerbucks 3. mystified- 4580 kerbucks 4. Johnster_Space_Program- 5250 kerbucks 5. DRAG0Nmon- 7,154 kerbucks Manned: 1. TheFlyingKerman- 2603 kerbucks 2. Aetharan- 3907 kerbucks 3. farmerben- 3982 kerbucks 4. Aetharan- 4276 kerbucks 5. Ironbars- 4423 kerbucks Notes- zolotiyeruki, if you can find some screenshots of that old craft, I will gladly accept it! Pds314 maintains the top 2 spots for lightest and cheapest unmanned crafts. TheFlyingKerman holds the cheapest manned rocket. Artharan has almost stolen every spot for the lightest manned rocket, taking 1, 2, 3, and 4! Will he be able to take 5? Keep up the amazing work everyone!

Ver 1.2 Adds all the parts for the 4 seat Altair Lander and AresV Launch Vehicle. This mod is to be used in conjuction with HOYO CSM mod and Ares 1-X mod. You have to download the Ares 1-X mod for the first stage Solid Rocket Boosters as well as 2nd stage(KDS) J2X engine and the HOYO CSM mod for the docking port. Place the Boosters on the AresV core stage fuel tank using the stock radial decouplers. You then should launch the Altair Lander and Kerbin Departure Stage(KDS) into orbit using the AresV. Launch the HOYO CSM using the Ares 1-X rocket to rendez-vous and dock with the Altair & KDS. Then use the J2X KDS engine to push the whole stack. Kerbal Joint Reinforcement is recommended so as to not have wobble problems due to the size of the launch vehicle parts and torque produced. If anyone has suggestions for weight/deltav etc. Please share them with me so i can improve the mod on the next update. Supports RealPlume, Engine Lighting and TAC Life Support. These mods are not bundled with the release but are highly recommended. Supports kOS and Telemachus. Also not bundled with the release. This mod comes bundled with dependencies. Module Manager, TexturesUnlimited (For Reflections), RasterPropMonitor (For functional interior), KSPWheel (For the landing legs. It is absolutely needed for the landing legs to work.), Starwaster's AnimatedDecouplers (For the decoupler animation), Alexustas' ASET Props 1.5 & ASET Avionics 2.1. INSTALLATION Be careful, this mod comes bundled with the latest version of ASET Props (1.5) & ASET Avionics (2.1). Delete these mods if you have older versions before installing this mod. Unzip and merge with your GameData folder. The folder GameData/LonesomeRobots from the zip must be merged with, not replace, any existing GameData/LonesomeRobots folder. DOWNLOAD https://spacedock.info/mod/1475/Altair Lander & AresV. LonesomeRobots Aerospace?ga=%3CGame+3102+%27Kerbal+Space+Program%27%3E .craft files for all LRAERO ships can be found here . These are saved from the latest KSP 1.3.1. Licensing. This mod is licensed under: Creative Commons Attribution-NonCommercial-ShareAlike 4.0 International Public License Redistributed with this mod. AnimatedDecouplers created by Starwaster. Redistributed as per license. RasterPropMonitor created by Mihara & MOARdV. Redistributed as per license. TexturesUnlimited created by Shadowmage. Redistributed as per license. KSPWheel created by Shadowmage. Redistributed as per license. ModuleManager created by Sarbian. Redistributed as per license. ASET Props 1.5 & ASET Avionics 2.1 created by Alexustas. Redistributed as per license. changes in this version Reflections are now handled by the awesome TexturesUnlimited mod. Texture Replacer is now deprecated. ModuleDeployableAntenna is now used to handle the Dish antenna animation. ---------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------- Ver 1.1 Adds all the parts for the 4 seat Altair Lander and AresV Launch Vehicle. This mod is to be used in conjuction with HOYO CSM mod and Ares 1-X mod. You have to download the Ares 1-X mod for the first stage Solid Rocket Boosters as well as 2nd stage(KDS) J2X engine and the HOYO CSM mod for the docking port. Place the Boosters on the AresV core stage fuel tank using the stock radial decouplers. You then should launch the Altair Lander and Kerbin Departure Stage(KDS) into orbit using the AresV. Launch the HOYO CSM using the Ares 1-X rocket to rendez-vous and dock with the Altair & KDS. Then use the J2X KDS engine to push the whole stack. Kerbal Joint Reinforcement is recommended so as to not have wobble problems due to the size of the launch vehicle parts and torque produced. If anyone has suggestions for weight/deltav etc. Please share them with me so i can improve the mod on the next update. Supports RealPlume, Engine Lighting and TAC Life Support. These mods are not bundled with the release but are highly recommended. Supports kOS and Telemachus. Also not bundled with the release. This mod comes bundled with dependencies. Module Manager, TextureReplacer (For Reflections), RasterPropMonitor (For functional interior), KSPWheel (For the landing legs. It is absolutely needed for the landing legs to work.), Starwaster's AnimatedDecouplers (For the decoupler animation), Alexustas' ASET Props 1.4 & ASET Avionics 2.0. INSTALLATION Unzip and merge with your GameData folder. The folder GameData/LonesomeRobots from the zip must be merged with, not replace, any existing GameData/LonesomeRobots folder. DOWNLOAD https://spacedock.info/mod/1475/Altair Lander & AresV. LonesomeRobots Aerospace?ga=%3CGame+3102+%27Kerbal+Space+Program%27%3E .craft files for all LRAERO ships can be found here . These are saved from the latest KSP 1.3.1. Licensing. This mod is licensed under: Creative Commons Attribution-NonCommercial-ShareAlike 4.0 International Public License Redistributed with this mod. AnimatedDecouplers created by Starwaster. Redistributed as per license. RasterPropMonitor created by Mihara & MOARdV. Redistributed as per license. TextureReplacer created by ducakar. Redistributed as per license. KSPWheel created by Shadowmage. Redistributed as per license. ModuleManager created by Sarbian. Redistributed as per license. ASET Props 1.4 & ASET Avionics 2.0 created by Alexustas. Redistributed as per license. changes in this version Changed ModuleRCS to ModuleRCSFX for all parts with RCS thrusters. Now when firing RCS produces sound. Landing gear now produce sound while retracting/extending. Changed TACLifeSupport configs. Now Altair lander has resources to support 4 kerbals for 9 days. 1 kerbal has resources for 24 days in the science lab. ---------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------- Ver 1.0 Adds all the parts for the 4 seat Altair Lander and AresV Launch Vehicle. This mod is to be used in conjuction with HOYO CSM mod and Ares 1-X mod. You have to download the Ares 1-X mod for the first stage Solid Rocket Boosters as well as 2nd stage(KDS) J2X engine and the HOYO CSM mod for the docking port. Place the Boosters on the AresV core stage fuel tank using the stock radial decouplers. You then should launch the Altair Lander and Kerbin Departure Stage(KDS) into orbit using the AresV. Launch the HOYO CSM using the Ares 1-X rocket to rendez-vous and dock with the Altair & KDS. Then use the J2X KDS engine to push the whole stack. Kerbal Joint Reinforcement is recommended so as to not have wobble problems due to the size of the launch vehicle parts and torque produced. If anyone has suggestions for weight/deltav etc. Please share them with me so i can improve the mod on the next update. Supports RealPlume and Engine Lighting. These mods are not bundled with the release but are highly recommended. Supports kOS and Telemachus. Also not bundled with the release. This mod comes bundled with dependencies. Module Manager, TextureReplacer (For Reflections), RasterPropMonitor (For functional interior), KSPWheel (For the landing legs. It is absolutely needed for the landing legs to work.), Starwaster's AnimatedDecouplers (For the decoupler animation), Alexustas' ASET Props 1.4 & ASET Avionics 2.0. INSTALLATION Unzip and merge with your GameData folder. The folder GameData/LonesomeRobots from the zip must be merged with, not replace, any existing GameData/LonesomeRobots folder. DOWNLOAD https://spacedock.info/mod/1475/Altair Lander & AresV. LonesomeRobots Aerospace?ga=%3CGame+3102+%27Kerbal+Space+Program%27%3E .craft files for all LRAERO ships can be found here . These are saved from the latest KSP 1.3.0. Licensing. This mod is licensed under: Creative Commons Attribution-NonCommercial-ShareAlike 4.0 International Public License Redistributed with this mod. AnimatedDecouplers created by Starwaster. Redistributed as per license. RasterPropMonitor created by Mihara & MOARdV. Redistributed as per license. TextureReplacer created by ducakar. Redistributed as per license. KSPWheel created by Shadowmage. Redistributed as per license. ModuleManager created by Sarbian. Redistributed as per license. ASET Props 1.4 & ASET Avionics 2.0 created by Alexustas. Redistributed as per license. SCREENSHOTS

This thread is sort of like the "Most Kerbal thing you have seen in real life" thread, just in KSP. It would probably have to be turning a Kania truck into a rocket. It was just for fun and I think I used infinite fuel to get it to space, but that's what the kerbals are about, right? Honestly, I haven't done that much kerbal-y stuff in KSP. I haven't really been to space that much. I'll put pictures of the truck here soon. EDIT: I just made a helicopter version that actually (pretty much) works. Now I am working on a truck/plane! ANOTHER EDIT: The plane worked ok, and I landed successfully, besides that I lost part of an engine during the landing. It only happened because I have the wheel brakes set too strong and the truck turned and rolled a little as it came to a stop.

Hi! I've got something new for y'all! Liquid Rhenium Solar Thermal Rocket The maximum temperature concentrated sunlight can heat a material to is 5800K. How do we approach this limit? We will describe existing and potential designs for solar thermal rockets. Solar thermal rockets The Solar Moth The principle of a solar thermal rocket is simple. You collect sunlight and focus it to heat a propellant headed for a nozzle. A rocket engine's performance is determined by its thrust, exhaust velocity and efficiency. A solar thermal rocket's thrust can be increased by sending more propellant through the nozzle. Its exhaust velocity can be increased by raising the propellant temperature. Doing either required more power, so more sunlight needs to be collected. Efficiency will depend on the design. The main advantages of a solar thermal rocket are its potential for high power density, high efficiency and high exhaust velocity. Collecting and heating with sunlight does not need massive equipment - unlike solar electric spacecraft that need solar panels, extremely lightweight reflective metal films can be used. A heat exchanger above a nozzle is compact and masses much less than the electrical equipment and electromagnetic or electrostatic accelerators a solar electric craft uses. Radiators are not needed either, as the propellant carries away the heat it absorbs with it. Put together, a solar thermal rocket can achieve power densities of 1MW/kg while solar electric craft struggle to rise above 1kW/kg. Sunlight would follow the same path as the laser beam here. As the sunlight is being absorbed by a propellant and expanded through a nozzle, there are only two energy conversion steps: sunlight to heat, then heat to kinetic energy. The first step can be assumed to be 99% efficient. The second step depends on nozzle design, but is generally better than 80%. Exhaust velocity will be determined by the root mean square velocity of the gas the propellant turns into. The equation is: Exhaust velocity: (3 * R * Temperature * 1000 / Molar mass ) ^ 0.5 Temperature is in Kelvins. Molar mass is the average g/mol value of the propellant at the temperature it is heated to. R is the molar gas constant, equal to 8.314 J/mol/K. For the very hot gasses we will be considering, we can assume complete dissociation of all molecules. H2 (2g/mol) will become atomic hydrogen (1g/mol), water (18g/mol) becomes a hydrogen-oxygen vapor (6g/mol) and so on. Low molar masses are preferred, with the best propellant being mono-atomic hydrogen unless other factors are considered. These advantages are all the critical elements that allow for travel throughout the inner solar system without requiring vast quantities of propellant. This means smaller spacecraft and lower travel times. Heat exchangers and exhaust velocity The limiting factor for solar thermal rockets is how hot they can heat the propellant. Directly heating the propellant is a difficult task. The lowest molar mass propellant, hydrogen, has terrible absorption. For all practical purposes, it is transparent to sunlight. Seeding the propellant with dust particles that absorb sunlight and heat the hydrogen indirectly through conduction has a major catch: the dust particles get dragged along by the hydrogen propellant flow and increase the average molar mass. A single millimeter-sized carbon dust particle in a cubic meter of hydrogen increases the molar mass from 1g/mol to Indirect heating involved using a heat exchanger as an intermediary between the sunlight collected and the propellant being heated. So far, designs have required the use of a solid mass of metal that is heated up by concentrated sunlight. The propellant is run over the metal, or through channels in the metal, to absorb the heat. Tungsten is often selected for this task, as it has a high resistance to heat, is strong even near its melting point and has a good thermal conductivity. Testing a Hafnium/Silicon Carbide coating. More modern designs make the most of the latest advances in materials technology to allow for higher operating temperatures. Carbon, notably, stays solid at temperatures as high as 4000K. Tantalum hafnium carbide and a new Hafnium-Nitrogen-Carbon compound melt at temperatures of 4200 and 4400K respectively. However, looking at our exhaust velocity equation, the limits of modern materials technology will only provide a 21% increase over common tungsten. This is the reason why so many propulsion technologies that rely on exchanging heat between a heat source, such as a nuclear fuel or a laser beam, and a propellant using a solid interface are said to be 'materials limited' to an exhaust velocity of 9.6km/s with tungsten, or 10km/s with carbon. THC or HNC would allow for an exhaust velocity of 10.5km/s. This is the deltaV equation, also known as the Tsiolkovsky rocket equation: DeltaV = ln (Wet mass / Dry mass) * Exhaust Velocity Wet mass is how much spaceship masses with a full load of propellant. Dry mass is the mass without any propellant. The wet to dry mass is also referred to as the 'mass ratio' of a rocket. We can rewrite the rocket equation to work out the required mass ratio to achieve a certain deltaV using a rocket engine's exhaust velocity: Mass ratio = e ^ (DeltaV required/Exhaust Velocity) 'e' is the exponent 2.7182... in simpler terms, the mass ratio increases exponentially as the deltaV required increases. Or, put another way, the mass ratio required decreases exponentially as the exhaust velocity rises. It is critical to have a higher exhaust velocity for rapid space travel without requiring massive rockets and towers of propellant. You might also have noticed that 'solid' is a keyword up to this point. Why must the heat exchanger remain solid? Liquid Rhenium There is a method to achieve the true maximal performance of a solar thermal rocket, which is heating up the propellant as far as it can go. This is incidentally the temperature of the surface of the sun (5800K). At this temperature, hydrogen propellant reaches an exhaust velocity of 12km/s. A rare, silver-black metal. Rhenium is a rare metal with a surprising number of qualities, one of which is a very high boiling point. Rhenium melts at 3459K but remains liquid up to 5903K. The trick to achieving higher exhaust velocities is to use a molten heat exchanger, specifically liquid rhenium at a temperature of 5800K. Rhenium is also very stable and does not react with hydrogen even at high temperatures, which is something carbon-based materials struggle to survive. It has already been considered as a heat exchanger, in solid form, by NASA. Here is a design that can use liquid rhenium as a heat exchanger: The diagram is for illustrative purposes only - a functional schematic would be more detailed. Here is an explanation for each component: Solar collector: A very large, very lightweight reflective film based on solar sails that can collect sunlight and focus it through a series of lens onto the heat exchanger fluid's inner surface. Rotating drum: The drum's inner surface contains a liquid heat exchanger. The outer surface is actively cooled. The drum is dotted with tiny channels that allow the propellant to enter the liquid from the bottom and bubble through to the top. It is made of Tantalum-Hafnium Carbide. Fluid surface: The fluid here is liquid rhenium. Its surface is heated to 5800K by concentrated sunlight. The lower layers nearer the drum holding the fluid is cooler. The centripetal forces hold the fluid in place Pressure chamber: The rotating gas mix gets separated here. Dense rhenium vapours fall back down, hot hydrogen escapes. Bubble-through heating: The rotation induces artificial gravity, allowing the hydrogen to heat up and rise through the denser rhenium. As it rises, it reaches hotter layers of the fluid heat exchanger. At the surface, it has reached 5800K. Small bubbles in direct contact with the rhenium allows for optimal thermal conductivity. More detail below. Active cooling loop: liquid hydrogen from the propellant tanks makes a first pass through the drum walls, lowering the temperature below the melting point of THC. It emerges as hot, high pressure gaseous hydrogen. High pressure loop: The heated hydrogen is forced through the channels in the drum. It emerges into the fluid heat exchanger as a series of tiny bubbles. Here is a close up of the drum wall, which contains both active cooling and high pressure channels: The configuration displayed above allows the hydrogen to enter the basin bottom at 4000K, then be heated further to 5800K before being ejected into the pressure chamber. If higher quantities of liquid hydrogen for active cooling are used, the drum and high pressure channel temperatures can be lowered to 3800, 3500, 3000K or lower. This pebble-bed nuclear thermal reactor has most of the components of our solar thermal rocket, except that instead using pebbles of nuclear, fuel, we use a liquid rhenium bed heated by sunlight. If the liquid hydrogen active cooling cannot handle the full heat load, radiators will be needed to cool down the drum below its melting point of 4215K. Thankfully, these radiators will receive coolant at 4000K. Their operating temperature will be incredibly high, allowing for tiny surface areas to reject tens of megawatts of waste heat. Electricity can also be generated by exploiting the temperature difference across the radiators' entrance and exit flows, and at very high efficiency. Operation The design is a Rotating Drum Fluid Heat Exchanger Solar Thermal Rocket (RD-FHE STR). It allows for hydrogen propellant to reach 5800K and achieve the maximum performance of a Solar Thermal Rocket. Liquid rhenium does not boil at 5800K, so it remain liquid and can be held inside the basin by simple centripetal forces. Vapor pressure of rhenium at 5800K (0.was determined to be low enough for our purposes. A surface of rhenium exposed to vacuum at that temperature would lose 0.076g/cm^2/s, or 762g/m^2/s. It is unknown how much centripetal force affects the loss rate of rhenium. The pressure chamber would operate at several dozens of atmospheres of pressure, which is known to increase the boiling point and reduce the evaporation rate of fluids. The same techniques used in Open-Cycle Gas Core nuclear reactors to prevent the loss of uranium gas can be applied to reducing the loss of rhenium vapours. At worst, the rhenium heat exchanger loses 0.76 kg of rhenium for square meter per second of operation. Looking at the designs below, the mass flow rate is measured in tons of hydrogen per second. This is a ratio of 1000:1, to be improved by various rhenium-retaining techniques. It should also be noted that rhenium is a very expensive material. A tungsten-rhenium mixhas very similar thermal properties and is much cheaper. Sunlight at 1AU provides 1367W/m^2. A broad-spectrum reflecting surface such as polished aluminium would capture and concentrate over 95% of this energy, so more than 1298W would be available per square meter. Solar sails materials such as 5um Mylar sheets are preferred, massing only 7g/m^2. More advanced materials technology, such as aluminium film resting on graphene foam, might mass as little as 0.1g/m^2. The 'Solar Moth' used inflatable support structure for its mirrors. Based on data for the Solar Moth concept, we have estimated that a solar thermal propulsion system can attain power densities of 1MW/kg. So, each square meter of collector area will require another 1.29 grams of equipment to convert sunlight into propulsive power. Performance Robot Asteroid Prospector We will calculate the performance of two versions of the RD-FHE STR. The first version uses modern materials and technologies, such as a 7g/m^2 Mylar sheet to collect sunlight and a 167kW/kg engine power density. The second version is more advanced, using 0.1g/m^2 sunlight collectors and a 1MW/kg power density. Modern RD-FHE 5 ton collection area => 714285m^2 927MW of sunlight focused onto the drum. 5.56 ton propulsion system Exhaust velocity: 12km/s Thrust: 123.4kN (80% efficiency) Thrust-to-weight ratio: 1.19 Overall power density: 87kW/kg Advanced RD-FHE 5 ton collection area =>50000000m^2 64.9GW of sunlight received 64.9 ton propulsion system Exhaust velocity: 12km/s Thrust: 10.8MN Thrust-to-weight ratio: 15.75 Overall power density: 928kW/kg The principal argument against solar thermal rockets, that their TWR is too low and their acceleration would take too long to justify the increase in Isp, can be beaten by using very high temperatures and very low mass sunlight collectors. For example, a 50 ton propulsion system based on the modern RD-FHE STR design, would be able to push 100 ton payloads to Mars (6km/s mission deltaV) using only 97 tons of propellant. It would leave Earth orbit at a decent 0.24g of acceleration, averaging 0.32g. The departure burn would take only 20 minutes. Using the advanced version of the RD-FHE solar thermal rocket would allow for a positively impressive acceleration of 3.1g. With 12km/s exhaust velocity, multiple missions that chemical rockets struggled to do with low-energy Hohmann transfers can be avoided. A chemical rocket such as SpaceX's BFR might achieve an Isp of 375s, which corresponds to an exhaust velocity of 3.67km/s. It would need a mass ratio of 5.13 to barely produce enough deltaV for a Mars mission. Earth to Destination. If our solar thermal rocket is granted the same mass ratio, it would have a deltaV of 19.6km/s. This allows for a Mars mission to be completed in under two months (10km/s departure, 9km/s insertion). It is also enough deltaV to reach Jupiter with a single stage. Other benefits include a vast reduction in the propellant-producing infrastructure needed to supply orbital refuelling depots and the ability to land on Mercury. Alternative versions: Blown hydrogen: Instead of bubbling hydrogen from the bottom of the liquid rhenium basin, hydrogen is blown into the pressure chamber from the top. It is heated by simply passing over the fluid heat exchanger. The advantage is that the rotating drum does not have to be riddled by microchannels, allowing it to be stronger and rotate faster, which would reduce rhenium losses, and also accept a higher rate of active cooling by leaving more room for liquid hydrogen channels. Another advantage is that there is less chance of hydrogen bubbles merging and exploding in showers at the surface, dragging along rhenium as they escape. The disadvantages is vastly reduced heat conduction rate between the rhenium and the hydrogen. This would require a long and thin pressure chamber to increase the time the hydrogen stays in contact with the rhenium, potentially making the propulsion system heavier than it needs to be and forcing sunlight to enter the chamber at very acute angles. ISRU propellants: Instead of hydrogen, other gaseous propellants might be used. Nitrogen is a good choice, as it is inert and only reduces the exhaust velocity by a factor 3.7 compared to hydrogen. Powering a hydrogen extraction process on Mars requires huge areas of solar panels. Nitrogen is easily sourced from Earth's atmosphere by gas scoops. Other options, such as water or carbon dioxide, are also viable and available on other planets. The advantage is that non-hydrogen propellants are easy to contain and are much denser than hydrogen, so their propellant tanks can be lightweight and small. They are easily sourced and only need to be scooped up and filtered, unlike hydrogen that has to undergo electrolysis. The disadvantage is that there propellants cannot serve as expandable coolant for the rotating drum. A radiator using a closed gas loop is necessary - helium is a likely candidate. This adds mass. A lower exhaust velocity also removes the principal advantage the RD-FHE STR has over other propulsion systems.

Ver 1.3 Launch vehicle for the HOYO CSM. Includes first stage solid rocket booster with 2 textures, J2X second stage engine, liquid fuel tank and decouplers/adapters. This mod is intended for the latest HOYO CSM (ver1.5) link at the end of the post. Supports RealPlume and Engine Lighting. These mods are not bundled with the release but are highly recommended. This mod comes bundled with dependencies. Module Manager, TexturesUnlimited (For Reflections) and FireSpitter (For SRB texture switching). INSTALLATION Unzip and merge with your GameData folder. The folder GameData/LonesomeRobots from the zip must be merged with, not replace, any existing GameData/LonesomeRobots folder. DOWNLOAD https://spacedock.info/mod/1314/ARES I-X. LonesomeRobots Aerospace Licensing. This mod is licensed under: Creative Commons Attribution-NonCommercial-ShareAlike 4.0 International Public License Redistributed with this mod. Firespitter created by snjo. Redistributed as per license. TexturesUnlimited created by Shadowmage. Redistributed as per license. ModuleManager created by Sarbian. Redistributed as per license. If you haven't tried the HOYO CSM yet check the following link http://forum.kerbalspaceprogram.com/index.php?/topic/154088-122113-lonesomerobots-aerospace-hoyo-csm .craft files for all LRAERO ships can be found here . These are saved from the latest KSP 1.3.1. ------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------ Ver 1.2 Launch vehicle for the HOYO CSM. Includes first stage solid rocket booster with 2 textures, J2X second stage engine, liquid fuel tank and decouplers/adapters. This mod is intended for the latest HOYO CSM (ver1.3) link at the end of the post. Supports RealPlume and Engine Lighting. These mods are not bundled with the release but are highly recommended. This mod comes bundled with dependencies. Module Manager, TextureReplacer (For Reflections) and FireSpitter (For SRB texture switching). INSTALLATION Unzip and merge with your GameData folder. The folder GameData/LonesomeRobots from the zip must be merged with, not replace, any existing GameData/LonesomeRobots folder. DOWNLOAD https://spacedock.info/mod/1314/ARES I-X. LonesomeRobots Aerospace Licensing. This mod is licensed under: Creative Commons Attribution-NonCommercial-ShareAlike 4.0 International Public License Redistributed with this mod. Firespitter created by snjo. Redistributed as per license. TextureReplacer created by ducakar. Redistributed as per license. ModuleManager created by Sarbian. Redistributed as per license. If you haven't tried the HOYO CSM yet check the following link http://forum.kerbalspaceprogram.com/index.php?/topic/154088-122113-lonesomerobots-aerospace-hoyo-csm .craft files for all LRAERO ships can be found here . These are saved from the latest KSP 1.3.1. ------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------ Ver 1.1 Launch vehicle for the HOYO CSM. Includes first stage solid rocket booster with 2 textures, J2X second stage engine, liquid fuel tank and decouplers/adapters. This mod is intended for the latest HOYO CSM (ver1.3) link at the end of the post. Supports RealPlume and Engine Lighting. These mods are not bundled with the release but are highly recommended. This mod comes bundled with dependencies. Module Manager, TextureReplacer (For Reflections) and FireSpitter (For SRB texture switching). INSTALLATION Unzip and merge with your GameData folder. The folder GameData/LonesomeRobots from the zip must be merged with, not replace, any existing GameData/LonesomeRobots folder. DOWNLOAD https://spacedock.info/mod/1314/ARES I-X. LonesomeRobots Aerospace Licensing. This mod is licensed under: Creative Commons Attribution-NonCommercial-ShareAlike 4.0 International Public License Redistributed with this mod. Firespitter created by snjo. Redistributed as per license. TextureReplacer created by ducakar. Redistributed as per license. ModuleManager created by Sarbian. Redistributed as per license. If you haven't tried the HOYO CSM yet check the following link http://forum.kerbalspaceprogram.com/index.php?/topic/154088-122113-lonesomerobots-aerospace-hoyo-csm .craft files for all LRAERO ships can be found here . These are saved from the latest KSP 1.3.0.

Remember this rocket from Captain Earth when Globe launched it to prevent the alien menace? You know, the "Kill-T-Gang"? Now, this rocket has become a reality, in 5 different versions! Ladies and Gentlemen, I (and Mitsubishi themselves) present to you, the Pegasus rocket family! The rockets are as follows: Pegasus (the original) Pegasus-E (for Expendable) Pegasus-L (for Liquid, hinting at the liquid-fueled boosters) Pegasus-P (from "poudre", French for solid), and Pegasus-L Hydrolox , a more powerful (and environmentally friendly) version of the Pegasus, with 4 Hydrolox engines. Also included is the "X Base One Propulsion Section" for "X Base One" (in turn both based on NASA's Constellation program and the BBC's Doctor Who (Bowie Base One, The Waters of Mars), and the Pegasus 2, Pegasus-M for the MRC, Pegasus-H for my original "Hippogriff" spacecraft, Pegasus-O for the Orion spacecraft, Unicorn, EuroPegasus, DORUS, and a version for the launch of Mir's core module. See them on my hangar right here: https://kerbalx.com/hangars/145355