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Found 49 results

  1. 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
  2. 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 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 .craft files for all LRAERO ships can be found here . These are saved from the latest KSP 1.3.0.
  3. Everything works fine until I try to make my gravity turn and it tilts north and goes out of control. Ive tried throttling down as I ascend but it still happens.
  4. Who here is excited for the brand-new Apollo Saturn V set that lego just released? Come on! I know that at least some of you are both Lego and KSP fans, so get talking! Who here is excited for this set, is getting this set, other stuff, etc. I basically just started this thread to discuss this new set, that was created by Lego fans Felix Stiessen and Valérie Roche, on the popular Lego Ideas website, where fans submit models they've made, and, if the model gets 10,000 supporters, Lego will review it, and, it it's passed the review process, make it into a lego set. So come on, get talking!
  5. Star Shooter KerbalX craft File here! So I made this rocket because I thought it would be a fun challenge and I've not really seen anyone do such a thing before. The first challenge came when I realized the "payload" would need to be placed near the bottom of the whole setup for purposes of easy attachment (I didn't want loading the Asteroid to be a complicated task). The 2nd hurdle was that the forward tank and engine nacelle would need to eject FORWARDS during flight and would need to ACCELERATE faster than the ships current speed to get out of the way (trust me it took way more tests to figure that out than probably should have!). The solution to this is to release the Front Tank/Engines when fuel is about to run out. The tank naturally accelerates forward and downwards safely clearing the path for the Star Shooter. During Stage 2 (tri-rocket configuration) you can fly to a stable orbit, or ditch the outer rockets so they can burn up in the atmosphere and continue your circulation burn on the Stage 3 dual Thud motors. I like to fly this rocket at half throttle until stage 2. After which full throttle is just fine. So far I've only tested it with a Class B Asteroid (technically meteorite) that weighed a mere 6.5 tons. I would be confident in payloads up to about 15 tons and of Class B asteroid size or smaller. Try it out and check out the video if you aren't convinced! Let me know what you think of her!
  6. Himalaya is a family of Kerbonian rockets named after the mountain of kerbal mythology, Himalæia. There are four known rockets, with two of them still being built. Some of them sent kerbals to space, others sent satellites to orbit, and others, well. you got me there. They were designed so that the U.S.K would beat the Koviet Union in the race of the kosmos. Himalaya II: The Himalaya II (KZ-59E) was the first of the Himalaya family to be manufactured. It has thirty parts, a height of 36.2 meters, a weight of 198,540 kg and a diameter of 3.75 meters. It costs 102,010, a TWR of 7.44 (max) and a Δv of 6,288 m/s. The length of the first mission was 1h 16m and 45s, and the crew was Jeb, Bill and Bob Kerman. Himalaya III: The Himalaya III (KZ-63B) is famous for launching the first satellite, GRAVIOLI (GRavity, Atmosphere and SurVey IOn-propelled SatelLIte), and was the second Himalaya manufactured. Information Unknown. Himalaya III-B Keverest: Still in construction. Himalaya IV: Still in construction.
  7. Post your Engineering Marvels! Since the genesis of mankind, the human race has observed growth of monumental proportions. Engineers and scientist, invented and inspired others, to create the best solutions, to all of man's problems. That is exactly what we are doing here! Below, you can show the world your amazing creations, an exhibition so to speak. This will show everyone your engineering talents, whilst also providing inspirations to others, to accelerate the KSP engineering !
  8. Welcome to my heavy lifter thread! So far I only have one, but more are planned. DOWNLOADS: DISCOVERY 1: SOON DISCOVERY I I made this craft waaaaay back in 1.0.5, yet it still remains my best orbital booster and Mün injection craft ever. Parts: 135 Mass: 432.024t TOTAL D/v: 2699 M/s Height: 41.4 meters. Payload to LKO: Unknown Uses for this craft: Fuel tug, Mün mission tug, orbital booster Screenshots:
  9. The thread where I will post all my designs and progress of my crafts! (With The Help Of Many Mods Of Course) ISE Dauntless [SSTO] Body, & propulsion systems will consist of parts from [Mark IV Spaceplane System] Wings, & lifting surface parts will consist of parts from [B9 Procedural Wings] So far.... Progress Report I Progress Report II
  10. Here I'll be posting my drawings of various spacecraft/rockets in MS Paint and where you can ask for drawings of other real-life rockets/spacecraft. I drew a Saturn V (image below) and thought, In the comments, tell me what I should draw next. (I'm already going to draw the Mercury Redstone, so don't suggest that.) It has to be a real-life spacecraft or rocket. Also, tell me what you think of them! They may not be the best in the world with detail (I actually forgot the ridges on a part of the Saturn V next to the right fin if you'll notice). Waiting List: 1. N1 - cubinator 2. ITS - TheEpicSquared 3. Gemini Titan - munlander1 4. Apollo LEM - Dafni 5. Apollo Little Joe - munlander1 6. Titan IIIe-Centaur - IncongruousGoat 7. DynaSoar - munlander1 8. Venture Star - Pine 9. Mercury Little Joe - munlander1 10. Open 11. Saturn 1b - munlander1 12. Open 13. V2 Bumper - munlander1 14. Open 15. Apollo 13 CSM (post-explosion) - munlander1 16. Open 17. Orion - munlander1 18. Open 19. CST-100 - munlander1 See ya!
  11. 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!
  12. Hello Guys, I´ve got a problem with my rockets. No clue why, but the tend to flip over after rolling a little bit. btw. I´m a beginner A time ago I didn´t have this problem, but with the time my rockets kept getting bigger and the problems started. That´s my rocket at the moment: Could you please help me prevent this or improve my rockets? Thank you IRobot
  13. Howdy, This craft sways once it achieves mid altitude. I was hoping that that someone fixes it while I go and eat dinner Any suggestions would be nice. Sigh, attachments seem to be very buggy... EDIT: Well I've fixed everything so now it is a working SSTO. Thanks, Chris (Heretic391)
  14. 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?
  15. I just want to share the design for a rocket I came up with that uses jet-fuel boosters to get its first kick to space rather than liquid fuel or SRBs. I haven't gotten this to space yet; I took a screencap shortly after (actual) liftoff and shortly after the game crashed. Image below" Honestly, I can't believe it even took off. It estimates about 18 minutes worth of fuel. This is a very early modded science mode game, and these are the first jet fuel tanks/engines/air intakes that you can unlock. The intakes and engines were made larger with Tweakscale, and that also increased the output enough for the bigger tanks. I haven't tried this with other engines. Has anyone else done this and, if so, how did it go?
  16. Decided to make an experiment and compile the data for something this basic, yet very important. So, here is it! Craft: stock "Jumping Flea" (Flea booster with MK1), straight launch, variable TWR (adjusted in VAB from data given by KER): TWR 1.2, highest real altitude: 4100 m flameout: 3100m TWR 1.5, highest real altitude: 6200 m flameout: 4300m TWR 1.7, highest real altitude: 7200 m flameout: 4600m TWR 2.0, highest real altitude: 8000 m flameout: 4700m TWR 2.3, highest real altitude: 8250 m flameout: 4700m TWR 2.5, highest real altitude: 8320 m flameout: 4600m TWR 2.7, highest real altitude: 8350 m flameout: 4500m TWR 3.0, highest real altitude: 8350 m flameout: 4200m TWR 3.3, highest real altitude: 8350 m flameout: 4000m TWR 7.6 (maximum), highest real altitude: 8100 m flameout: 2100m Explaination: this means that: at TWR lower than 2, your craft will be burning fuel fighting the gravity, and at TWR higher than 2.5, your craft will be burning fuel fighting atmospheric drag. Enjoy! ## Update 2016/09/30 Quick check with Swivel+FL-T400+2xz1k batts+RC001s+Adv.NoseconeA (2,156 DeltaV) confirms the data: TWR: 1.7, highest real altitude: 87,600 m flameout: 31,000m TWR: 2.2, highest real altitude: 117,500 m flameout: 31,000m TWR 2.2: ## Update 2016/09/30 - 2 In response to criticism, I decided to make a small "usable" 7 tone "payload" (Payload A). The payload has everything to make orbit, keep in orbit and descend down. To repeat, this thread about ideal TWR for first stage / launching stage ### Payload A, first test Payload A config: Protective Nose Cone Mk7 + LCR01 RGU + Srvc.Bay 2.5m(3x Z1k batt, 2x SP-L solar panels, 2x MK2-R radial chutes) + Advanced Reaction Wheels, Large + X200-8 Fuel + LV-909 "Terrier"+Rockomax Decoupler. --- Payload A weight: 7,000 kg This payload will be launched using only one stage - using two ascend methods, using two different TWRs. The methods: 1) direct climb for statistics: SAS on, max throttle, ignition. No Stage 2 disconnect. 2) manual gravity turn: - 0 degree until 2,5km; - course to 15 degree(75 on navball) at 2,5km; - course to 35 degree (55 on navball) on pass 15km; - course to 55 degree (35 on navball) at 25km; - course to 75 degree (15 on navball) at 35 km; (this step was skipped on 2.2 TWR due to atmospheric heat **) - deactivate if apoapsis reaches 80km, wait until T-45s (accordingly, this step was much longer on 2.2 TWR) - burn at horizon line under prograde till end of stage: output stage 1 data - disconnect stage 1 and finalize burn using stage 2: output payload A in orbit fuel left will be also noted: out of fuel/flameout for 1st stage (altitude at; horizontal and vertical speeds), amount of fuel left on payload after circularization at 80/80km. #### Payload A (stage 2) - uplifter A (stage 1) uplifter A config: (Payload A +) Jumbo 64 tank + RE-M3 Mainsail (+3x launch stability enhancer, start only) 3,218 projected stage 1 DeltaV; 6,370 projected total DeltaV ) [ craft file link ] Results: ##### Direct climb data TWR 1.25, highest real altitude: 268,000 m flameout: 72,000m TWR 2.20, highest real altitude: 615,000 m flameout: 69,000m ##### Manual gravity turn data info: TWR 2.20 appears to compress air higher (hidden aerodynamic center) - thus rocket becomes less stable and either smoother transition between nodes (used here) or additional lifting surfaces required. TWR 1.25: apoapsis at flameout: 82,000m periapsis at flameout: -390,000m Stage 2 80km/80km orbit fuel left: fuel - 247/360; LOX - 302/440 TWR 2.20: apoapsis at flameout: 86,000m periapsis at flameout: -22,000m Stage 2 80km/80km (90/90km actually) orbit fuel left: fuel - 340/360; LOX - 416/440 ** apparently there is a room for improvement! "Engine idling on suborbit ascend equals wasted fuel". This case is free to be tested! Probable method: throttle down after leaving atmosphere (~28km). Reason: it looks like with TWR 2.2 atmosphere heat prevents efficient acceleration, requiring higher angle of attack. But according to Oberth effect, lower attitude accelerations are more efficient, thus its better to take it slower below.
  17. So i have been fascinated with the concept of a valveless pulse-jet pretty much since i learned of its existence. Largely due to its simplicity. I've had the thought before that as a inexpensive engine it might work as a 1st stage for a rocket. It's air breathing and as such doesn't need complex oxidizer setups. I made one in KSP with reasonably accurate TWR and a few other things, and mentioned it in my project thread but got very little response (so it probably lacks some accuracy). Question is, what could be done (if anything) to improove its thrust/efficency? I know there was a bit of development in the 60's but it more or less started and ended there. This probably isn't terribly scientific but, any ideas?
  18. Please help! Every single rocket i build slowly slides off to one side, no matter what. Everything is in line, center of mass is in the bottom middle if my rocket, i am using fins, sas is on, and im not touching the engine. I have no idea what is going on and how to fix it. The game is nearly unplayable because i can't reach orbit to collect science so i can't unlock halfway decent parts. Please help
  19. Hey there intelligent Engineers! I took a look at the latest rocket designs and realised, that most of them are using kerosine for first stage, which does emit a hell lot of CO2. Then I thought about making a rocket ECO-friendly, and started looking for a propellant, which would fit my needs. And voila! LH2 and LOX or better known as Liquid Hydrogen and Liquid Oxygen. These fluids put together produce H2O, which is very eco-friendly. Luckily LH2 and LOX engines have the highest specific Impulse, and therefore the highest efficiency. Now I challenge you guys, to build a entirely Eco-friendly Rocket, which can carry a payload into LEO (in RSS ofc.)atleast! The rules are: Eco Firendly fuel, like LH2 and LOX No SRB's No kerosine or any kind of enviromentaly harming/unfriendly rocket fuel. RSS (Real Solar System) Rocket has to achieve a minimum orbit of 180 km. As an example I built an eco-friendly rocket called ECO-1, which can carry around 20t to LEO (approx. 200 km). For the engines I used 6x SSME ( Space Shuttle Main Engine; Aerojet Rocketdyne RS-25 ). Here are some pictures of it: Im excited to see your designs, rockets etc. Thanks for your attention!
  20. Does anyone know an easy way to make a rocket overheat? I need it to for a YouTube cinema type video I'm making.
  21. Hello! Welcome to my hanger of the crafts I made. As by the title these will be replicas. I will be doing planes and rockets. They are all stock craft. These wont be perfect but I'm just trying my best so please give me some helpful criticism. I hope you like these very much! More planes will come soon! List of Planes Lockheed L-1011 "Tristar" DHC-8 "Dash-8" MD-80 "Mad Dog" Cessna 172 Cessna Grand Caravan EX Airbus A340-600 Cessna Citation Mustang Saab 340 List of Rockets V-2 Proton M Titan IIIE "Centaur Planes Lockheed L-1011 "Tristar" The Lockheed L-1011 was a plane made by the Lockheed Corporation in 1970 and entering service in April of 1972. They built 250 aircraft and the plane is still in limited use today. It had 3 engines, 2 under the wing and 1 in the tail using a S-duct design. The plane does handle slowly but it still can fly. Download see Lockheed L-1011 "TriStar" on DHC-8 "Dash-8" The Dash 8 was a plane made by De Havilland Canada and Bombardier Aerospace in 1983 and introduced in 1984. As of December 31,2015 1,179 planes have been built. It is still in production and service today. Download MD-80 "Mad Dog" The MD-80 was A plane devopled by McDonnell Douglas in 1979 and introduced in 1980.There were 1,191 aircraft built and still in service. Download see MD-80 "Mad Dog" on Cessna 172 The Cessna 172 is the most successful plane in history with over 43,000 planes built and longevity. The cessna was frist made in 1955 with inroduction in 1956 with many variations over the years. (NOTE) Please switch to the front cockpit in flight. Download see Cessna 172 on Cessna Grand Caravan EX The Cessna Grand Caravan was a plane developed by Cessna 1982 and intorduced in 1984. Cessna has built 2,500 planes and still in production. It’s still in service today primarily as a cargo aircrat. (NOTE) Be carefull on landing or the gear will explode. Download see Cessna Grand Caravan EX on Airbus A340-600 The Airbus A340 was developed by Airbus in Mid 2001 and it was introduced in Mid 2002. 97 of this variant was built, this aircraft is still in service. Download see Airbus A340-600 on Cessna Citation Mustang The Mustang developed by Cessna in Mid 2005 and introduced in Late 2006. Cessna has built 425 so far. The plane is still in service. Download see Cessna Citation Mustang on Saab 340 The Sab 340 was developed by Saab and Fairchildwith its early first flight in 1983 and intoduced in 1983. 459 planes were built and it’s still in service. Download see SAAB 340 on Rockets V-2 The V-2 was a missle devopled by the Nazis around 1942. it produced from 1942-1945 but some were built after the war. It’s no longer in use. Download see V-2 on Proton M The Proton M was a rocket mde by Khrunichev around 2001 and the first flight was in April of 2001. There have been 98 launches with 10 failures. The rocket is still active today.¨This rocket has been tested with a 85T paypoad to a 100 by 100 KM orbit. Download see Proton M on Titan IIIE "Centaur" The Titan IIIE was a rocket made by Marlin Marietta in 1967 and it’s first use in February of 1974 It was flow 7 times with 1 failure. The rocket is no longer in use.¨This rocket has beeen tested with a 54.5T payload in a 100 by 100 KM orbit. Download see Titan IIIE "Centaur" on
  22. HELLO ALL This "1 Station 1 Rocket" Idea is essentially the concept of using 1 booster stage that you may not recover by going into the VAB or SPH again. LEADERBOARD: to put a space station into an orbit of any altitude. Station MUST be 3 modules or more, each module must (tbd) contain it's own probe, batteries, it's own way to generate power and must be 1.25m and be a fair size, just... Honor code people, no cheeky .625m Oscar B sized modules. STEPS: - Build a booster stage to get the payload to a point where it can use it's onboard engines to boost it to orbit - Build a fuel truck to refuel the rocket after landing the booster - Build a crane to put the module on the rocket Put the module on the booster however necessary - Pack enough fuel in the modules to assemble a space station using only one rocket ---------------------------------------------------------------------------------------------------------------------------------------- HARD MODE: 1 Moon base 1 Rocket STEPS: - Same thing as station except put it on the moon ---------------------------------------------------------------------------------------------------------------------------------------- SCORING: I will keep a leaderboard of the scores, essentially you will be rated as to the mass of the station and rocket (KGs) minus the cost in funds (station and rocket), just so we have a cheap but large space station. (If you have a better idea plz comment)
  23. Good news everyone! Do you have a hankerin for the good old days? Well now you can fly like they wanted to in the early 1960's Introducing the Farnsworth Dyna-soar. 1.1.3 friendly Back in the early 60's Boeing designed this neat rocket concept. Unfortunately, it never left the ground. Instead of the glider being attached to the side of the rocket like we are accustomed to seeing with the space shuttle, Boeing attached it to the top of the rocket. The space shuttle suffered from some inherent design flaws right off of the drawing board (due mainly to government budget cuts). The main one being that it has the thrust and mass are out of alignment. To compensate for this the designers mounted the SSME's (Space Shuttle Main Engines) at an angle so that the thrust could be brought inline with the center of mass. Unfortunately, this flight attitude brought the orbiter directly under the external tank. The orbiter was pretty fragile and could not handle much fod (foreign object debris)hitting it. The external tank was huge, covered in finicky insulation and was directly over the orbiter during assent. A great way for the orbiter to get hit with stuff... The External tank that was filled with liquid oxygen and liquid hydrogen. Those tanks had lines on them while it was on the pad to boil off some of those gasses that were in liquid form to control the pressure in the tanks. The result was the external tank would get cold enough sometimes for patches of ice to form on the tank out of condensing humidity in the air. Technicians would check for ice before launch but invariably, some would be missed. At launch by the time the shuttle cleared the launch tower it was already hurtling towards space at over 200 mph. Plenty fast enough for a piece of ice mingled with some foam from the external tank to break off and slice through the brittle carbon/carbon leading edge of the orbiter wing. This is the best guess as to what happened to the Columbia in 2003 ...Sorry for the history lesson... Anyways, The Dyna-Soar avoided all of these problems by mounding the glider on the top of the rocket. This however limited the size of the glider. It required the modified Titan rocket to have large fins on the bottom to compensate for the glider moving the center of lift too far up the rocket. So here is my take on the Dyna Soar All stages can take advantage of the "stage recovery" mod if you use it. The glider retains the last stage to use for orbital maneuvers, docking and solar energy. 150k orbits are achievable for crew rotation at your station. Final stage has a Clamp-0-Tron jr docking port to attach to the station. More fun than exploding! Perform your deorbit burn and then detach from the final stage. The glider should enter the athmosphere perpendicular and with the elevators deployed. There is no fuel you can move around. Once the air thickens up, the glider will drop to about 30 deg. attitude. The glider is not a "true" glider. It has two juno engines and 100 units of fuel so if you over or under-shoot your landing there is some adjustments that can be made. The glider flies very well and has an awesome glide slope. This does make it hard to burn off airspeed. Try to cross over the runway at well under 100 m/s and you shouldn't overshoot. Download here!
  24. With the same 5150 delta-V you've come to expect, here is the Cyclone Heavy. Capable of reaching Duna directly with some clever aerobraking, the C-H can lift payloads of up to 20 tons to a 100 km orbit and return. It is fully reusable, except for the possible exception of the in-fairing docking port. It is reasonably forgiving, but don't do anything stupid with it. Do not open the fairing at altitudes lower than 55000 meters. Doing so will void your warranty. The C-H can be fitted with a hitchhiker container for crew storage (By default it is unmanned). It has no action groups or solar panels to worry about; the Cyclone Heavy is powered by 4 Radiothermal Isotope Generators and thus does not require external power input. It weighs an astounding 773 tons in the default configuration, with a sea level TWR of 1.58. If you so desire, there is more info is in the KerbalX page. Fly with care, or with mechjeb. It's stable at 3x time warp, and does not fall apart at 4x, but I advise you to stick to low timewarp speeds. (From left to right: Misttrailer SSTO, Cloudstream SSTO, Cyclone Heavy SSTO)
  25. BackStory time: The KSC has almost run out of funds but because they were focused on other planets the didin't check the mun or minmus. Challenge: Get to the mun or minmus (which you prefer) flyby or orbit or landing with the least expensive rocket you can Rules: No hyper edit. Can Be Kerbaled. Needs pics or video (i lost my proof in freak accident with hard drive) NO PART MODS. Mech jeb and KER but MJ for readouts only. Sandbox mode. Scoring: Funds divided in half and if gone to minmus x2 and return x #(if flyby then 1 if orbit then 2 if landing and return with kerbals 4) = Score ScoreBoards: 1: @Reactordrone 66215 points / 3107 2:@Signo 1700 points / 4800