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sevenperforce

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Posts posted by sevenperforce

  1. 33 minutes ago, magnemoe said:

    I see it likely they use some insulation between the bottom heat shield and the equipment there as the steel shield can manage much higher temperature than the equipment. For the sides this might be an issue, the cone reduce the  heating but you still get some radiation heating. 

    By the time they are doing re-entry, the tanks are over 90% empty. Autogenous-press gas is a relatively good insulator. At the most they might need a little something around the base to deal with radiation heating, but even there the emissivity of stainless steel is probably working in their favor.

    Have we ever seen a cutaway of their tank layout? If insulation does become an issue, I could see them doing concentric tanks or some kind of conical arrangement. Perhaps this:

    stoked.png

    Grey is the engine/heat shield section, while blue is a toroidal LOX tank and yellow is the LH2 tank. This way, the LOX tank does double-duty as double-wall insulation for the hydrogen tank during re-entry, because the remaining liquid hydrogen will be well below the single-walled region. Of course then you have to deal with insulation on the inner common bulkhead between the LOX and the hydrogen.

  2. 16 hours ago, Arugela said:

    86400tons*2000=172800000*47.880172=8273693721.6/2205=3752242.05061224489795918367

    What measurement unit is it?

    Or is that the wrong math?

    Should it be 8273693721.6/2000=4136846.8608*2205= 9121747328.064

    Are any of these an an existing unit?

    This is the wrong math. You're not providing any units at all (except for "tons" up at the top, and even then I don't know whether you're talking about metric tonnes (1000 kg) or Imperial tons  (2000 lbs).

    Include all your units for the entire equation. Units are algebraic; treat them like variables and multiply/add/divide as necessary. Without including units, your numbers are gibberish.

    Also, can you contextualize the system a little bit for us? I have no idea what you're trying to figure out, which makes this all very challenging to characterize. Is this a launch vehicle? Are you burning in a straight line or along a curved trajectory? Are you already in orbit when the burn begins? All of this is important.

    16 hours ago, Arugela said:

    Does that mean that the two formulas added together is simply the complete distance traveled?

    No. Use one formula at a time. If you need to combine formulae, use unit substitution (e.g. if you know that a = b + c2, but c = b-3, then you can substitute b-3 in place of c to get a = b + (b-3)2 so as to solve for in the equation).

    16 hours ago, Arugela said:

    And what measurement unit is that then? My brain goes screwy still when going between lb and metric.

    Best way to avoid your brain going screwy over units is to actually use units consistently. 

    Metric units, unlike Imperial units, have the auspicious property of automatically cancelling if you use the correct base units. For example, if I know the propellant consumption (kilograms/second) and effective exhaust velocity (meters/second) of a jet airliner traveling at cruising speed, I know I can multiply them together to get cruising thrust, because (kilograms * meters) / (second * second) equals kg*m/s2 which is equal to Newtons. And since Newtons of thrust are equal to Newtons of drag at constant speed, I can divide those Newtons by ρ*CD*(velocity)2/2 (where ρ is air density and CD is the drag coefficient) to get the cross-sectional area of the aircraft, because Newtons are kg*m/s2, and kg*m/s2 divided by density (kg/m3) and velocity squared (m2/s2) gives me units of m2 which is area.

    In those situations I may not even bother writing out my units because I know they are going to cancel out properly. But since you're not familiar with the correct base units I strongly suggest keeping track of the units consistently and doing the cancellations manually, or you'll miss something.

    Obligatory funny:

    Spoiler

    dimensional_analysis.png

     

  3. 11 hours ago, magnemoe said:

    Downside is that some might point an telescope at the satellite, GEO satellites was pretty rare in the 70's, some astronomer might point an telescope at it. Also you now only see one part of earth. 

    If it's farther than GEO then it will still be able to see all of Earth over the course of its orbit.

    Yes, some amateur astronomer might point a scope at it, but it would be difficult to pick up again if it was in an orbit like that. Also limited reflectivity would make it virtually invisible. 

    11 hours ago, magnemoe said:

    Now being responsible for various religious phenomena is weird, could be done with some sort of reentery probe with holograms but here I say magic.  
    Still trying to board it to loot the alien technology would be very tempting. 

    I don't see how it could be responsible for anything on Earth but then again I haven't read the book.

  4. 12 hours ago, SunlitZelkova said:

    https://x.com/nukestrat/status/1699443848538190175?s=46&t=Jd73T2beq0JLNtwTy1uR5A
     

    Wow. This wasn’t any old Minuteman test. It was a test of the MIRV option, with three warheads!

    Currently under treaty regulations, each missile only carries one warhead. But due to current events…

    A sign of the times to come?

    My understanding of New START was that MIRVs are fine, but each warhead counts as an individual missile even when clustered on a single missile for the purposes of deployment limits. But maybe I am incorrect.

  5. 4 hours ago, darthgently said:
    7 hours ago, RyanRising said:

    I wonder how they're handling insulation. It must be inside the outer shell because we can see bare steel there, so something similar to S-IVB maybe?

    It is cooled internally by cryogenic fuel flow much like an engine bell from what I recall.

    That's how they are cooling the heat shield, but insulation is a different issue entirely -- keeping the propellant from boiling off due to ambient heat transferred through the skin.

    To @RyanRising's question -- my guess is that these early hopper prototypes have simply dispensed with insulation altogether. The low fineness of the stage minimizes surface area, anyway. For short hops and even suborbital hops, the amount of propellant lost to boiloff is going to be relatively negligible.

    Once they have an actual orbital vehicle under construction, we'll see whether they do some sort of insulation. It may be that the additional weight of insulation would be greater than the weight of propellant lost to boiloff anyway, and so eschewing insulation entirely is the best approach.

  6. 29 minutes ago, Arugela said:

    I'll assume that has to do with the relative inefficiency of rockets? I assume because of the excessive thrust needed.

    No, that doesn't have anything to do with inefficiencies; it's a fundamental aspect of rocket engines. A rocket engine has to push something out of the back end in order to produce forward thrust.

    29 minutes ago, Arugela said:

    You are always technically throwing fuel out something... Do batteries loose weight also? Not sure how electric works.

    Well, some batteries lose weight during discharge and others gain weight during discharge, but most remain about the same. However, that's not quite the right comparison. You can use up a battery in any number of ways on board a spacecraft, but unless you're yeeting some sort of exhaust out of the back of your spacecraft, your spacecraft isn't going to move. Electric ion thrusters operate by using electrical energy to accelerate tiny amounts of ionized gas at high velocity, but even they are subject to changes in acceleration as the gas is used up and the vehicle becomes lighter. The Dawn spacecraft, for example, carried a total of 425 kilograms of xenon gas, about 35% of its total weight at launch.

    29 minutes ago, Arugela said:

    Yea, I was trying to figure out distance over delta V. I was having a hard time finding how to figure that out. If constant fuel use is a thing can't you just add to the acceleration from the fuel mass loss?

    If fuel use is constant then yes, you have a constant change in acceleration and you can determine the distance covered accordingly. You just have to go to a third-order kinematic equation instead of the more familiar second-order kinematic equations.

    Velocity is the rate at which position changes; an object moving at 10 m/s is changing its position by 10 meters every second. Acceleration is the rate at which velocity changes; an object accelerating at 10 (m/s)/s or 10 m/s2 is changing its velocity by 10 m/s every second. The rate of change in acceleration is called "jerk" (or sometimes "jolt") and it has units of ((m/s)/s)/s or m/s3

    The third-order kinematic equation for distance can be readily derived by integration, but I'll skip that step and just give it to you:

    x = x0 + v0t + 0.5*a0t2 + 1/6jt3

    In order to solve this equation, you'll need to know each of the starting values (x, v0, and a0) as well as your value for jerk (j). In your case, the starting position x seems like it is going to be zero, and the same appears to be true for the starting velocity v0 as well. So you need to define a0 and j.

    If your initial acceleration is 7.3392 m/s2, then that's your a0. How do we calculate j? Well, with constant propellant use, j will be a constant, so we can use this equation (should be self-evident; let me know if it's not):

    j = (af - a0)/t

    In order to accelerate a starting mass of 86,400 tonnes at 7.3392 m/s2, you'll need an engine producing 634 meganewtons. By the end of your burn, your vehicle mass has dropped to 21,600 tonnes but you're still thrusting at 634 meganewtons, giving you a final acceleration of 29.36 m/s2. Using our equation above, we get jerk equal to 29.36 m/s2 minus 7.3392 m/s2 divided by 600 seconds, or 0.0367 m/s3.

    Now that we have that, we plug all the rest in. 

    x = x0 + v0t + 0.5*a0t2 + 1/6jt3

     

    x = 0 + 0*600 sec + 0.5*7.3392 m/s2 * (600 sec)2 + 1/6*(0.0367 m/s3)*(600 sec)3

    x = 0 + 0*600 sec + 3.6696 m/s2 * 360,000 s2 + 0.00612 m/s* 216,000,000 s3

    x = 1,321,056 m/s2 * s2 + 1,321,920 m/s* s3

    x = 2,642,976 meters

    So there's your answer -- which you will note is about double the answer you would have gotten without accounting for change in acceleration.

    29 minutes ago, Arugela said:

    I'll assume that is the formula for rho etx and air density mixed in.

    Nope, it is not. Kinematic equations assume no air resistance. If you're talking about something moving through an atmosphere then we're gonna have a whole other set of issues.

    45 minutes ago, Arugela said:

    We really need standard formulas for acceleration in imperial as well as metric.. It gets confusing.

    The standard kinematic formulae are unit-agnostic.

  7. I've been out of pocket. Wenhop?

    11 hours ago, darthgently said:

    I wonder how much real time control (aside from flight termination trigger) they will have over the craft at various points in the ascent.  There is probably onboard code to automatically attempt a target suborbital apoapsis and eccentricity with a ceiling on reentry speed and reentry geographic location in mind, but can they intervene remotely and raise periapsis at apoapsis instead?  I'd love to know all the details. 

    Fairly certain that everything after about T-30 seconds is almost completely automated. They program in all of the parameters and tell it what to do with all contingencies and give it abort modes for anything out of spec.

  8. 3 hours ago, RCgothic said:

    You appear to be calculating s=0.5*a*t^2, which is the formula for distance traveled under constant acceleration starting from rest over a time interval.

    @Shpaget has the correct formula for velocity under constant acceleration starting from rest over a time interval.

    As usual @RCgothic is entirely correct but I will also point out, just in case OP missed it, that the constant acceleration equation will not work properly for a reaction engine. If you are shoving exhaust out of a nozzle, then you also have to account for the decrease in your total vehicle mass across the period of the burn. You'd need to use the Tsiokolvsky rocket equation for that.

    15 hours ago, Arugela said:

    Actually I need to make sure 600 isn't the result of the 600 seconds. If so then change 600 to time or time+1. If it's a seperate value then keep it the same.

    The 600 is most definitely the result of the 600 seconds. You're just doing iteratively what s=0.5*a*t^2 does non-iteratively.

  9. 35 minutes ago, tater said:

    Possibly dumb question: The hiring laws are in fact separate from the ITAR  (and EAR?) laws, so what happens with possible goal conflicts? Ie: if you think that security risk/liability trumps hiring liability.

    The hiring laws prohibit unlawful employment discrimination (on the basis of race, color, religion, sex, or national origin) but the ITAR laws trump them, because refusing to hire someone based on ITAR is not unlawful. In other words, you aren't refusing based on national origin; you're refusing based on ITAR status. So if you are a foreign national who applies to work for ULA in an ITAR-sensitive role, ULA can honestly say "Sorry, we can't consider your application, at least not without a specific ITAR exemption from the State Department, because ITAR will not allow us to share information with you."

    Where SpaceX screwed up (intentionally or unintentionally) was saying the same thing to refugees and asylees with legal status even though ITAR does not prohibit them from hiring refugees and asylees. 

    To add to my prior examples: it would be like a hospital telling a gay man, "We can't hire you because the FDA prohibits gay men from coming into contact with blood" when in fact the FDA rule (which has now been rescinded, finally) prohibited gay men from donating blood but said nothing about healthcare employment generally. If it was JUST a job posting, that might be chalked up to a mistake, but if the hospital was also refusing to consider applications from gay men who applied for maintenance work or billing jobs (e.g., no physical patient contact) then it would look much more like a deliberate lie to disguise homophobic discrimination.

  10. 2 minutes ago, tater said:
    1 hour ago, sevenperforce said:

    Then I used WayBack to look up historical postings and found this:

    Oh, so they need to change the specific laws cited on their job postings?

    Thanks for that analysis, makes more sense now.

    Yeah, basically. But not just the laws they cite -- it's also the terms they use, and the terms they don't use.

    It's like if the admissions office at a university had a posting that said "Federal law prohibits us from considering student applications without proof of Selective Service registration or proof of exemption." Since only males can register or be exempted from Selective Service, this would imply that federal law prohibits the university from accepting female students, which obviously is not true. Such a posting might just be a mistake/misstatement, but if there was evidence that the contractor refused to accept female students on this basis (e.g., if it refused applications from women on the basis that they did not provide proof of Selective Service registration) then the DOJ could take enforcement action under Title IX.

    Or imagine a job posting for a teacher at an elementary school which said "Bright or exotic dyed hair can be a distraction to students: acceptable hair colors for applicants include blonde, strawberry blonde, platinum, light brown, and auburn." This suggests that individuals with naturally black hair (e.g., people of a particular marginalized ethnicity) need not apply. Again, this could just be a miscommunication or mistake, but if they actually turned away applicants based on hair color then they could get in trouble.

  11. 1 hour ago, tater said:

    DOJ can say whatever they like, but the actual legality is for the courts to decide, not DOJ—so a suit is maybe the best solution to clarify things. 

    The DOJ lawsuit sets out some pretty specific claims.  Among these is the claim that SpaceX used its public announcements, job applications, and online recruiting communications to exclude asylees and refugees. This isn't hard to verify: the internet is forever, after all.

    Then I used WayBack to look up historical postings and found this:

    • To conform to U.S. Government space technology export regulations, including the International Traffic in Arms Regulations (ITAR) you must be a U.S. citizen, lawful permanent resident of the U.S., protected individual as defined by 8 U.S.C. 1324b(a)(3), or eligible to obtain the required authorizations from the U.S. Department of State. Learn more about the ITAR here.  

    And then I looked up a current job posting on the SpaceX recruiting website (emphasis shows additions):

    • To conform to U.S. Government export regulations, applicant must be a (i) U.S. citizen or national, (ii) U.S. lawful, permanent resident (aka green card holder), (iii) Refugee under 8 U.S.C. § 1157, or (iv) Asylee under 8 U.S.C. § 1158, or be eligible to obtain the required authorizations from the U.S. Department of State. Learn more about the ITAR here.

    So that's the difference. It might not seem like a huge difference, but it's a meaningful one: as job postings go it plainly violates Title VII, by discouraging applicants based on national origin.  An inaccurate job posting by itself isn't proof of discriminatory intent, but taken together with the other evidence proffered by DOJ's lawsuit it tends to support their allegations of broadly discriminatory practices (particularly if the DOJ notified them and they didn't correct it).

    1 hour ago, tater said:

    One law saying people can be held criminally liable for tech transfer, another saying you must consider hiring randos who walked across the border and claimed asylum (there's supposed to be a paperwork process before coming over, but that is not always the case).

    Form I-589 (the paperwork required for application for asylum) can be filed before crossing the border or arriving at a point of entry to seek asylum, but it does not have to be filed in advance. Federal law gives asylum-seekers up to a year after they enter the country before they are required to file (18 USC § 1158(a)(2)(B)). In fact, if initial screening is conducted by USCIS, then USCIS will automatically treat the screening itself as a filing of the application.

    However, an "Asylee under 8 U.S.C. § 1158" would be a person who has not only filed Form I-589 but has actually been granted asylum under subsection (b)(1)(A) of the code section. So, not some rando who walked across the border and just claimed asylum.

  12. 3 hours ago, Spacescifi said:

    Every nozzle has a rocket engine behind it typically.

    If it's a working engine, yes.

    3 hours ago, Spacescifi said:

    You could actually make rocket engine feed multiple nozzles

    This is very common in Russian designs and other engines with Soviet legacy, but less common in Western designs (the Rocketdyne XLR-89-5 is a notable exception). However, it's important not to confuse the entire engine assembly with individual combustion chambers. Here's an example, shown by the venerable RD-180:

    engines-vs-chambers.png

    This is one liquid propellant rocket engine with a single thrust structure, a single gimbal mount, a single turbopump, and a single propellant flow system. However, there are two combustion chambers and two nozzles.

    I suspect that when you are calling something an "engine" you are actually talking about the combustion chamber. You would never want to make a single combustion chamber feed multiple nozzles, because the pressure loss along the path from the chamber to the various nozzles would be too inefficient. The only exception would be for a very low-impulse system, like a set of attitude control thrusters or propulsive vents.

    Try to think about individual rocket engine combustion chambers like the individual pistons in an automobile engine and the entire rocket engine assembly like the completed automobile engine+fuel pump+crankshaft system.

    3 hours ago, Spacescifi said:

    but I think it is more common if you need good thrust to use mutiple engines. Since one engine has a limit on power, which can be increased with more engines of similar or greater power.

    There is no real limit on the power from a single engine. When you're talking about rocket engines, you need to be conscious of both power input and power output. The power input is the amount of power that is generated by the turbopump in order to push the propellants into the combustion chamber(s). The turbopump is basically a jet turbine engine all on its own and can be almost arbitrarily large; the fuel turbopump for the Aerojet M-1 engine produced a 75,000 horsepower, and the RD-170's single-shaft turbopump produced a whopping 230,000 horsepower. That's up there with the horsepower of the largest diesel engines ever produced, although it is dwarfed by the effective horsepower of airliner engines (like the GE-90's 21 trillion horsepower).

    The drive to multiple combustion chambers was not an issue of power, but an issue of combustion stability. Once you end up with a gigantic combustion chamber (the volume of the F-1 engine combustion chamber was almost a full cubic meter), the flow of massive amounts of propellant starts to produce currents that can impede combustion. That's why the Russians used multiple combustion chambers, and it's one of the reasons that a company like SpaceX would rather have dozens of smaller Raptor engines than build a gigantic Super Raptor.

    3 hours ago, Spacescifi said:

    Main Question: What happens if your chamber pressure is too high and your nozzle is too small? I have been told that bigger nozzles are needed for higher chamber pressures, but was not told the reason why.

    Lemme guess {boom}.

    Nope, not boom. Who told you that bigger nozzles are needed for higher chamber pressures? Or did you just assume that? That's not how it works.

    When exhaust comes yeeting out of the combustion chamber, it has a lot of heat and a lot of pressure. The nozzle allows the exhaust to expand, trading heat for velocity and providing a surface against which the expanding exhaust can push to produce thrust. Complete expansion is impossible without an infinitely long nozzle, so you have to truncate the nozzle somewhere.

    Pressure is the force that a fluid exerts on a containing surface (and on itself). Force is proportional to acceleration, so the higher your chamber pressure, the more rapidly it will accelerate as it travels down the nozzle. At very high combustion pressures, you can get away with a relatively short nozzle, because of how quickly the exhaust expands. At low combustion pressures, you need a longer nozzle to give the exhaust time to fully expand. So it's exactly the opposite of what you were told.

    If your nozzle is not large/long enough for your design pressure, it won't hurt anything; you're just wasting pressure that could otherwise be converted into thrust (the exhaust will spill out the edges of the nozzle before it has fully expanded).

    3 hours ago, Spacescifi said:

    Secondary Question: Higher chamber pressure equals more thrust right? But there is a limit on how high the chamber pressure can be too right?

    Higher chamber pressure equals more thrust for an equivalent sized nozzle. If you increase the pressure, the forces acting on the nozzle are greater, and so even though you still have spillover at the edge of the nozzle, you will have gotten more out of the exhaust flow by the time it reaches the edge of the nozzle.

    If you want to increase thrust directly, you have to increase the amount of propellant going into your engine, usually by spinning the turbopumps faster. There is a limit to how fast you can spin your turbopumps before you just kinda have to build a bigger turbopump.

    There is no real limit on how high the chamber pressure can be, except for the fact that a higher chamber pressure requires a heavier chamber (to contain the pressure) and a beefier turbopump (to push the propellant into the chamber at a higher pressure) which means your engine will be much heavier overall. So there's a tradeoff between chamber pressure and engine weight. There's also a tradeoff between chamber pressure and turbopump power consumption. The turbopump typically burns a portion of your propellant to provide the power to push the rest of the propellant into the chamber, but if you ramp up your chamber pressure too high, then you end up using up too much of your propellant inside the turbopump.

    3 hours ago, Spacescifi said:

    Third Question: Can you have multiple rocket engines on ONE nozzle? Or is that just silly and no good reason for it because you will probably melt the nozzle trying that?

    It's perfectly fine to have multiple combustion chambers feeding a single nozzle -- it's common in aerospike designs. Certainly no problem with melting the nozzle, if the nozzle is regeneratively cooled.

    I'm not sure why you're always afraid of things melting. Melting is certainly a common challenge in the world of rocket science, but it's a technical/efficiency challenge, not a basic limitation of physics. Melting is usually NOT the limiting factor on making an engine bigger or more powerful.

  13. On 8/26/2023 at 4:26 PM, Exoscientist said:

    First, make a smaller rocket with, say, 9 Raptors... Test the heck out this vehicle at full up, full thrust, and full flight duration static tests. Just they like did and still do for the Falcon 9...

    SpaceX does not conduct "full flight duration" static fire tests of any of its launch vehicles. When SpaceX announces that it has completed a "full duration" static fire, it means that the static fire went the full duration it was intended to go. This is rarely more than a few seconds; the longest Falcon Heavy static fire went for a dozen seconds. AFAIK, no launch vehicle in history has ever conducted a full thrust, full flight duration static fire test. The closest analogue is probably the Green Run testing campaign for SLS which sometimes goes full flight duration but of course is only operating at a little over 20% of full liftoff thrust.

    On 8/26/2023 at 4:26 PM, Exoscientist said:

    Only after multiple full static tests where all engines successfully pass...

    ...something that has never been done on any rocket, ever...

    On 8/26/2023 at 4:26 PM, Exoscientist said:

    ...do they send this up for a test flight.

    What, precisely, is the advantage of this imagined full-thrust full-flight-duration static fire over a test launch?

    If your goal is to line the pockets of whatever company is supplying parts for ultra beefy hold-down clamps, I suppose this makes sense. No other reason though.

    On 8/26/2023 at 4:26 PM, Exoscientist said:

    Such a two-stage rocket could do 100 tons to LEO. That is sufficient to do single-launch lunar and Mars missions. Now, make money on that rocket made from the start to be reusable.

    How do you figure?

    A booster with a third of the liftoff thrust of Superheavy cannot possibly deliver 100 tonnes to LEO with full reusability when Superheavy+Starship can only do 120-150 tonnes to LEO with reusability.

    If you mean flying smaller payloads with full reusability, then it becomes really questionable whether you're reaching any advantages over Falcon 9 and Falcon Heavy, since the reuse of your "mini-Starship" becomes challenging. How do you EDL the upper stage? You can't just scale down the current Starship design, because even the "skydiver" entry works the same way, your landing won't: Raptor Vacuum is vastly overpowered for a hoverslam landing. 

    If you mean smaller payloads with only first-stage reuse, then you really have no advantage over the Falcon family.

    On 8/26/2023 at 4:26 PM, Exoscientist said:

    Fly that rocket very many times all the while making profit with it.

    SpaceX already has the market cornered with Falcon 9 alone. Why do you think they aren't flying Falcon Heavy more often? The answer is simple: when Falcon 9 can deliver virtually every commercial payload with first stage reuse, there's really no need to add more cores.

    On 8/26/2023 at 4:26 PM, Exoscientist said:

    THEN after flying so many times...form your superheavy lift vehicle by using triple cores, like what happened with the Falcon Heavy.

    Going from Falcon 9 to Falcon Heavy was a nightmarish dev path. The Falcon Heavy core is a completely new rocket compared to the Falcon 9 booster. If you want a Raptor-based vehicle with more than 9 engines, it's a better plan to just go big from the start.

    On 8/26/2023 at 4:26 PM, Exoscientist said:

    you’ll be making money on the smaller version and be able even to make both single-launch Moon and Mars missions with it

    A scaled-up Falcon 9 upper stage powered by a single Raptor Vacuum would suffer from high dry mass, making it much less useful for single-launch missions BLEO.

    On 8/26/2023 at 5:58 PM, mikegarrison said:
    On 8/26/2023 at 4:30 AM, RCgothic said:

    The new foundation isn't going to fail again. It's over two meters thick reinforced concrete capped with steel and with many more support pilings than previously.

    That's what they thought the first time....

    Unless I misremember, SpaceX was very much anticipating that substantial pad damage would be a potential outcome. They thought there was a chance it might survive, but they knew there was a chance that it wouldn't, and figured that as long as they got good launch data, it would be worth it.

    They didn't anticipate the degree of pad destruction, but clearly it wasn't catastrophic given that they have already gotten it functional again.

    On 8/27/2023 at 10:24 AM, Exoscientist said:

    I am dismayed by the level of unreliability of the Raptor.

    Do you have any data to support your dismay?

    On 8/27/2023 at 10:24 AM, Exoscientist said:

    In this latest static fire test two of the Raptors had to be shut  down for only a 5 second test and 50% thrust level.

    Do you have any data that suggest the early shutdowns were an engine reliability issue?

    On 8/27/2023 at 10:24 AM, Exoscientist said:

    I don’t think anyone doubts that if there is another test launch in like two weeks there will be engine failures like before.

    I don't think anyone outside of SpaceX and perhaps certain people at the FAA know whether there were actually any engine failures in the first flight or not.

    There were engine shutdowns, yes. There are no data that suggest one way or another whether these engine shutdowns were commanded.

    22 hours ago, .50calBMG said:

    How many RS-25s blew up during testing over how many years before becoming operational? IIRC only one RS-25 ever failed after that...

    While I think we are generally in agreement, I will note that RS-25 failures have been responsible for a total of seven launch aborts: six with the Shuttle and one with SLS. In one case the RS-25 failure was seconds away from causing LOCV.

    This just goes for the proposition that engine-out capability is generally a very good thing.

  14. 17 hours ago, Exoscientist said:

    The discussion was about using the SLS for cargo. I think it is too expensive for that purpose. Use the Falcon Heavy for that purpose, or other low cost commercial launchers.

    Oh, I certainly agree with that. SLS is far too expensive for cargo. Far too expensive for anything, really, but definitely for cargo.

    17 hours ago, Exoscientist said:

    I was talking about using two hydrolox stages to get the max cargo on the Falcon Heavy but here’s another way using a single stage:

    Suppose we use a 45 ton prop load “Centaur-like” stage carried to LEO by the Falcon Heavy for cargo only transport to the lunar surface one-way.

    If it's "Centaur-like" then you're not getting 45 tonnes of hydrolox inside even the Falcon Heavy extended fairing, as I explained above (and re-explain in more detail below).

    17 hours ago, Exoscientist said:

     As it’s Centaur-like, take the ISP as 465.5s(the max Centaur RL-10 Isp with extended nozzle was 465.5s)

    The RL10-C-2-1 has never been paired with a Centaur, but is still in production and achieves your desired specific impulse. It's a whopping 163.5" long but fortunately 77.2" of that is a deployable nozzle extension, leaving the 64.1" engine and the 22.2" fixed nozzle extension to give a total stowed height of 86.3". 

    Here are the internal dimensions of that extended fairing, lifted straight from the F9 user guide. The numbers on the left are height above the payload adapter; the numbers on the right are diameter:

    Spoiler

    Falcon-Extended-Fairing.png

    I'm not sure what cargo you're thinking of, but for the sake of simplicity let's cut the hydrolox stage off at ST = 477.976", right where the ogive starts. That means your payload, whatever it is, needs to fit inside a right truncated cone with a height of 175", a base with diameter 180", and a top with diameter 49". Let's allow 4" of clearance all the way around the stage for downcomers, vents, and the like (remember that this thing has to perform some significant maneuvering), giving us a stage diameter of 172". Tank wall thickness is of course negligible. Rule of thumb is that the most efficient dome shape is an ellipsoid of height R/2(1/2), so this gives an ellipsoidal cap height of 60.8". The double-walled intermediate bulkhead has a thickness of 0.3". Neglecting any volume occupied by stringers and the like, the total available propellant volume is represented as the sum of the volumes of an oblate spheroid of height 121.6" and a cylinder of height 269.8", both with a diameter of 175":

    Falcon-Fairing-with-Centaur-Plus.png

    The ellipsoid has a volume of 31.95 cubic meters and the cylinder has a volume of 106.34 cubic meters, giving this "Centaur Plus" frankenstage a total available internal propellant volume of 138.29 cubic meters. At hydrolox's bulk density of 0.28 g/cc, this gives 38.7 tonnes of propellant.

    Now, the RL10-C-2 has a propellant mass flow rate of 24.07 kg/s, giving it a total burn time of 1,608 seconds. That's far, far too long -- a full third of an LEO orbit (although obviously it wouldn't happen all at once; I'm just illustrating) -- so the Oberth losses would be immense. You're going to need two engines. It will be a tight squeeze under the payload fairing but it can probably be done.

    17 hours ago, Exoscientist said:

    ...and give it a ca. 10 to 1 mass ratio...

    If we're going "Centaur-like" then let's apply the Centaur mass ratio closely. Centaur has an empty mass of 2,247 kg, which drops to 2,032 kg when you subtract the mass of the RL10-C-1. It carries a total of 20,830 kg of hydrolox, giving it an engineless mass ratio of 10.25:1. Thus to hold 38.7 tonnes of propellant, our frankenstage needs a dry mass of 3,775 kg. Add the increased weight of two RL10-C-2s and the stage mass comes up to 4,377 kg.

    17 hours ago, Exoscientist said:

    Then with 12 tons payload...

    With 12 tonnes payload, the combined stack develops 5,540 m/s of Δv.

    Falcon Heavy can deliver 63.8 tonnes of payload to LEO, but that necessarily includes its own residuals. The PAF can't handle a ~55 tonne payload so it would need a special stage adapter to brace against the bulkhead or something. Estimating 1.8 tonnes for that structure, Falcon Heavy's "payload" is 56.9 tonnes and it reaches LEO with 6.9 tonnes of its own propellant to spare. Estimating FHUS at 4.6 tonnes, this means FH can burn the last of its props to give the upper stage stack thingy a boost of around 363 m/s past LEO, which is at least something helpful.

    Thus the total Δv available to the payload from LEO is 5.9 km/s. TLI is 3.2 km/s, LOI is 0.9 km/s, and descent to the lunar surface from orbit is 1.87 km/s, for a total requirement of 5.97 km/s. Just shy.

    However, you do have to actually LAND your payload, and our Centaur Plus frankenstage isn't going to be useful for that. If you go with a crasher stage architecture, then we can easily give the actual payload some small pressure-fed hypergolic thrusters to perform the last 70 m/s of the landing burn, plus whatever is necessary for the actual landing maneuvers.

    Of course, a 12-tonne payload is enough for a lunar ascent vehicle. On the other hand, this would require building a completely new intermediate Centaur frankenstage. IIRC, Centaur V was announced ca. 2012 and they still haven't ironed out the kinks, so we can assume a similar timeframe for this sort of retrofit.

    If you want to just play rocket legos with the existing Centaur SEC, that's much more doable. With the same intended 12-tonne payload, the existing Centaur SEC develops 3,982 m/s of dV and has a stack mass of 35.1 tonnes. It fits easily in the extended fairing with significantly more room for payload. Launched on Falcon Heavy (using a notional 1.3-tonne payload adapter frame), the Falcon upper stage would reach LEO with an impressive 27.4 tonnes of propellant residuals, allowing it to deliver a heft 1,747 m/s to the Centaur and its payload. This means the whole stack boasts 5.7 km/s of Δv, just 200 m/s less than the first design despite using only about half as much hydrolox. AND that's with the lower specific impulse of the RL10-C-1.

    Just goes to show that specific impulse ain't everything. Once you are going beyond LEO, mass ratio becomes more important than specific impulse, which is why the Merlin 1DV and the single RL10-C-1 can beat a pair or RL10-C-2s. The high specific impulse of hydrolox is more useful for lifting large monolithic payloads into LEO in the first place (think: Saturn V second stage).

  15. 18 hours ago, magnemoe said:

    I say modifying FH would be much easier than making an new lunar lander from scratch. Length of upper stage is based on RP1 an hydrolox state could be far longer

    F9 is already close to the maximum fineness ratio for a rocket. Wind conditions at altitude are a problem with high fineness because the bending moment increases dramatically. A significant tank stretch isn't really on the table.

    11 hours ago, Exoscientist said:
    On 8/21/2023 at 11:46 AM, sevenperforce said:

    With a hydrolox upper stage, a fully expendable Falcon Heavy can only send 9.3 tonnes of payload to TLI, let alone to the lunar surface. Hydrolox is fluffy; a hydrolox Falcon upper stage would only be able to carry around 30 tonnes of propellant.

     I’ve seen numbers for Falcon Heavy to TLI in the range of  ~20 tons

    Oh, absolutely. FH can absolutely send around that much to TLI. Which is why replacing the upper stage with hydrolox for only ~9-10 tonnes to TLI wouldn't be a good idea.

    11 hours ago, Exoscientist said:

    But to maximize payload don’t use the kerolox FH upper stage to do the TLI burn. Use the 63.8 ton FH capacity to LEO to carry hydrolox stages for the TLI  burn and for the lunar lander stage. For example a 30 ton Centaur-like stage with a 10 ton Centaur-like stage could get about 15 tons one-way to the lunar surface.

    I'm confused -- are you now talking about a four stage vehicle? Or 4.5 stage, counting the FH side boosters?

    Look at pages 83 and 88 of the Falcon 9 User Guide. The extended fairing has a cylindrical payload area 478" high and 181" in diameter, and there's an additional 175" of conical payload volume above that cylindrical plane. Let's imagine that your notional TLI hydrolox stage used a cluster of BE-7 engines to maximize utilization of volume and needed about 4" for tank walls, external fittings, and clearance from the fairing. The BE-7 engines are 80" tall. Let's imagine a lunar lander slightly squattier than the Apollo LM, at 5 meters height. That cuts 23" into the cylindrical region. Finally, let's assume a 1" insulated common bulkhead and shave off another 18" equivalent of vertical volume to account for the ellipsoidal caps.

    So that leaves a total available tank volume of 137 cubic meters, which gives us space for 38 tonnes of hydrolox. That's best-case-scenario assumptions with a single-stage architecture. If you try to stack two hydrolox stages on top of each other, your total combined propellant load drops to 28 tonnes of hydrolox.

    I'm still not quite sure what exactly you're proposing in terms of the mission profile, but from the overall context I think you are envisioning a lander+stage combo being launched on Falcon Heavy, meeting up with separately-launched crew somewhere in cislunar space, and then taking the crew down to the lunar surface and back up. For safety and simplicity, I'll consider a crasher stage architecture, where a zero-boiloff hydrolox stage performs both the braking burn into lunar orbit as well as the descent burn, and drops off to allow the lander to perform the final hovering landing as well as the ultimate ascent. Let's use the 9-12 tonne ascent vehicle concept from the Artemis initial studies; I'll go with 10 tonnes to make sure we have enough extra mass for landing legs and hovering propellant.

    The crasher stage needs 2.77 km/s of Δv to brake into cislunar space and then take the lander down to the surface. Assuming 450 seconds of specific impulse, the stack propellant fraction needs to be on the order of 47%. Assuming a relatively decent stage mass ratio of 10:1 including engine(s) and insulation system, the stage needs to be 52% of the total stack mass, giving us a total stack mass of 20.8 tonnes.

    But here we see there's no need for stacking up multiple hydrolox stages. Falcon Heavy can already deliver ~20 tonnes to TLI, so adding an additional stage underneath is completely unnecessary.

  16. 15 hours ago, Exoscientist said:

    For in-space hydrolox stages on the Falcon Heavy, I estimate 15 tons one-way to the lunar surface if you have low-boiloff tech. If only the stage for TLI is hydrolox, so you don’t need low boiloff, and with a storable propellant lander stage I estimate 10 tons one-way.

    With a hydrolox upper stage, a fully expendable Falcon Heavy can only send 9.3 tonnes of payload to TLI, let alone to the lunar surface. Hydrolox is fluffy; a hydrolox Falcon upper stage would only be able to carry around 30 tonnes of propellant.

  17. 17 minutes ago, Exoscientist said:
    4 hours ago, darthgently said:

    Ok, but as excellently pointed out by sevenpercentforce they can't travel backwards up the plume

    Sorry, but this is incorrect. Shockwaves or any kind of wave can be reflected any direction.

    Shockwaves cannot travel upstream in a supersonic flow. If they could, then supersonic jets would be destroyed by shockwave reflections.

    But more to the point, it's now become unclear what you're talking about here, exactly. Are you saying that the exhaust plumes could be reflected off of the water deluge and impinge on the booster? Or are you saying that the sonic shockwaves produced by engine sound pressure could be reflected off of the water deluge and cause more harm to the booster than sonic reflection off bare concrete?

    If the former, the answer is simply no. Exhaust plumes don't bounce.

    If the latter, what you're suggesting is at least physically possible, but since the water is not in a steady-state flow that is both planar and laminar, it's not possible here.

    17 minutes ago, Exoscientist said:

    This is a key point. Take a look at this video by Everyday Astronaut:

    <snip>

     The reverberating shockwaves visible at the 1:57 point in the @Erdayastronaut video suggest the SpaceX approach to water deluge can damage the engines.

    The fact that rocket engines inevitably produce reverberating sound waves -- something we have known for a long long time -- does not lead to the unsupported conclusion that this specific water deluge system would cause damage to the engines. There is neither a mechanism of action for such damage nor evidence that such a mechanism is in play.

    17 minutes ago, Exoscientist said:

    SpaceX should stop dismissing the lessons of Apollo and learn from them. Use a flame trench.

    Which Apollo test taught the "lesson" of using a flame trench? The Apollo launch site was built from the ground up with a flame trench. So you're not really advocating that we learn from "the lessons of Apollo"; you're just saying "they should do it like Apollo did it" without any reference to lessons.

    The pads that launched the Saturn V never had a sound suppression deluge at all; their meager water spray was for cooling and fire prevention. A sound suppression deluge was not added until after the first few Shuttle flights, to reduce sonic shockwave tile damage.

    And SpaceX does have a flame trench. It has six of them, in fact. The Saturn V's flame trench was 13 meters deep; the Orbital Launch Platform for Superheavy is 25 meters deep. 

  18. 8 hours ago, Exoscientist said:
    10 hours ago, sevenperforce said:
    16 hours ago, Exoscientist said:

    Then depending on the pressure of the water at this height, the exhaust flow impinging on it can cause reverberating pressure waves back on the engines.

    Exhaust flow is supersonic, so pressure waves can't travel backwards.

    Explosion shockwaves for example can bounce off of atmosphere layers and bounce back to Earth known as atmospheric focusing:

    Atmospheric focusing is a type of wave interaction causing shock waves to affect areas at a greater distance than otherwise expected. Variations in the atmosphere create distortions in the wavefront by refracting a segment, allowing it to converge at certain points and constructively interfere. In the case of destructive shock waves, this may result in areas of damage far beyond the theoretical extent of its blast effect. Examples of this are seen during supersonic booms, large extraterrestrial impacts from objects like meteors, and nuclear explosions.

    Atmospheric focusing can take place when a shockwave interacts with a discrete atmospheric boundary later, similar to the way that light will be partially refracted and partially reflected at the boundary layer between water and air or between glass and water. This requires, however, that the boundary layer exist in the same medium through which the wave is propagating. Shock waves are pressure waves that travel through the air, and since the air is the thing that has the boundary layer, the boundary layer can refract the shock wave.

    The exhaust flow from the business end of a Raptor engine is not a pressure wave in a static medium; it's a supersonic flow of exhaust. Supersonic flows are not shock waves, and the mass of chaotically-moving water spray is not a propagation medium. 

    Even if this was an apt analogy, which it isn't, a pressure wave experiencing atmospheric focusing is still going to be traveling away from the source, not back toward it.

    8 hours ago, Exoscientist said:

    When impinging on another object liquid, gas, or solid they can travel in any direction, like a ball headed backwards after hitting a wall.

    When a near-laminar supersonic flow impinges on a surface, it doesn't bounce. Rather, the collimation is destroyed and the flow experiences a transition normal to the original direction of travel, carrying the energy away in the normal plane.

    Sound pressure is an issue, of course, but the water significantly damps the sound pressure. There's no reflection off the water spray.

  19. 5 hours ago, Exoscientist said:

     This @Erdayastronaut clip shows quite alot of water reaching the level of the engines even though the water is angled outwards at the base.

    Water, like everything else, travels in a parabolic arc. So if it is angled outward at the base, it will be traveling at a shallower angle once it reaches the level of the engines, meaning that it will miss the engines.

    5 hours ago, Exoscientist said:

    Then depending on the pressure of the water at this height, the exhaust flow impinging on it can cause reverberating pressure waves back on the engines.

    Exhaust flow is supersonic, so pressure waves can't travel backwards.

    5 hours ago, Exoscientist said:

     In the 5+ second Booster 7 static fire in February, only two engines failed. In the test flight in April with a longer time on the pad to liftoff with a veritable concrete tornado throwing up chunck’s of concrete, only 3 engines failed in the initial liftoff. In this latest test, 4 engines failed after only 2.7 seconds.
    Cause:  the water deluge.

    Purely speculative and highly implausible.

    Also, a commanded shutdown is not the same as a failure. CRS-1 had an engine failure on ascent. A commanded shutdown during a static fire doesn't tell us anything.

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