Author Topic: New better engines & new material discovery may allow for space plane/SSTO  (Read 107917 times)

Offline john smith 19

  • Senior Member
  • *****
  • Posts: 10351
  • Everyplaceelse
  • Liked: 2431
  • Likes Given: 13606
Back in the 90's, there were just as many people who thought we could pull it off, if not more.  REL, for example, were pitching the same ideas for Skylon in the 90's.  And big bucks were being invested in NASP and X-33/VentureStar.
NASP would not have been funded if it had had an external audit before it was funded. When it did the the project was shut down, as TA Heppenheimer's "Facing the Heat Barrier" records.

  X33 failed for a number of reasons, mostly a combination of NASA wanting the maximum amount of new technology on top of a very demanding core task coupled with LM apparently using it at as a training exercise IE not putting their "A" team on the job.

Only SABRESkylon has survived a full external audit of the engine and structural plans by ESA and scrutiny at an international colloquim of launch vehicle and hypersonics researchers.

You keep comparing a partly funded design against 2 failed projects (implying it is also a failure).
This is a tactic common to Marketing departments running a negative campaign against a competitor.

In reality SABRESkylon is the only credible design still standing from the 3 listed.
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline john smith 19

  • Senior Member
  • *****
  • Posts: 10351
  • Everyplaceelse
  • Liked: 2431
  • Likes Given: 13606
I like this approach a lot, except that you run into extreme flow separation if you try to use that same engine for vertical landing.
There is an implicit assumption that the kerosene has all been burnt on the launch. The obvious answer is to retain some for the landing burn as well. The down side is you either reduce the payload by that amount or the whole vehicle gets proportionately bigger.

Whenever people talk about launch vehicles and engine engineering there are a lot of implicit assumptions.  You should be very wary of any system being ruled in or out without those assumptions being spelled out.

Some are correct, some were correct and some are just flat out wrong, such as the notion a leak in a LOX cooled combustion chamber would result in near instant failure.  The last has been known to be a gross exaggeration for over 20 years, yet so many engineers still treat it as a taboo.  :(
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Only SABRESkylon has survived a full external audit of the engine and structural plans by ESA and scrutiny at an international colloquim of launch vehicle and hypersonics researchers.

None of which have ever built a launch vehicle.
Human spaceflight is basically just LARPing now.

Offline nec207

  • Full Member
  • ***
  • Posts: 303
  • Liked: 3
  • Likes Given: 2

  Well exactly the 90's SSTO engines where not powerful enough to take it up 10 feet off the ground that alone take it up into space!!

There was big holes with all the 90's SSTO programs.  Not one of them shown that an engines could take it up into space. 

They put way too much money into 90's SSTO programs and the technology was just not there.


Wrong.  It had nothing to with engines.  There is no difference between engines then, now or even in the 60's/70's.  The issue is not the engines.

There were many engines that could get into space much more than 10 feet  The issues was getting into orbit with payload.

Can you elaborate on this? More so the one in bold?

If you want to learn the why's involved, I suggest you take a look at Kirk Sorensen's excellent series of articles here: http://selenianboondocks.com/category/rocket-design-theory/page/2/  , starting with this one: http://selenianboondocks.com/2010/02/rocket-equation-mod-1/

The relevant part for "getting into orbit" would be this one: http://selenianboondocks.com/2010/03/payload-fraction/

and this for an example of how more Isp is not necessarily better for getting into orbit if T/W is sacrificed: http://selenianboondocks.com/2010/06/ssto-ntr-bad/


The problem with the X-33 and lot of other 90's SSTO and space planes was heat tolerance in the aerospike engines. Composite and ceramic materials weren't far enough along to stand up to the higher than standard temperatures.

Ironically a few months after Bush slashed NASA's budget down to almost nothing, new materials solved the heat problem!!!  :o :o


Offline Asteroza

  • Senior Member
  • *****
  • Posts: 2836
  • Liked: 1084
  • Likes Given: 33
What materials in particular solved the aerospike heat problem? (I'm assuming throat/chamber heat here?)

Offline Eric Hedman

  • Senior Member
  • *****
  • Posts: 2314
  • The birthplace of the solid body electric guitar
  • Liked: 1953
  • Likes Given: 1144
The problem with the X-33 and lot of other 90's SSTO and space planes was heat tolerance in the aerospike engines. Composite and ceramic materials weren't far enough along to stand up to the higher than standard temperatures.

Ironically a few months after Bush slashed NASA's budget down to almost nothing, new materials solved the heat problem!!!  :o :o

Here is NASA's budget history.  If you look from fiscal year 2001 to 2009 the budgets for NASA in 2014 dollars (budgets proposed by Bush 43) went from 18.94 Billion to 19.714 Billion.  Congress pretty much set the final budgets, so when did Bush slash NASA's budget?  I also thought the big issue of many still unresolved was the delamination of the composite tanks.

https://en.wikipedia.org/wiki/Budget_of_NASA#Annual_budget.2C_1958-2015

Offline Nilof

  • Full Member
  • ****
  • Posts: 1177
  • Liked: 597
  • Likes Given: 707
Dry mass is the key source of payload uncertainty, which has been a much bigger obstacle for SSTO project funding than payload (i.e. independent engineers saying "this won't reach orbit with payload" andh aving a good argument).

But no independent engineers are actually saying that in this case, at least not ones that have verifiably looked at the design for more than five seconds.  The pessimism tends to exhibit a distinct lack of specifics, and to me it sounds more like a mixture of NIH and sour grapes left over from the '90s than any sort of actual well-grounded judgment.

You're essentially dismissing a high-resolution preliminary design (Skylon D) built on an ESA-vetted detailed concept (Skylon C) with 30 years of risk reduction and design iteration behind it starting from an attempt to address the problems uncovered by a thorough, well-funded design study (HOTOL) on the basis of a crude personal rule of thumb.

Some SSTO concepts amount to not much more than a couple of papers full of inconsistent, obviously fudged mass properties and engine performance parameters and handwaving about "next-generation TPS".  Skylon is not one of those.  REL have stated that the purpose of Skylon D is to anchor the engine design - the idea is to have enough confidence that the vehicle can be built to spec that the engine can be sized and developed first.  This implies a level of understanding of the mass budget that is much better than you are apparently assuming, and suggests that requiring 100% mass growth margin at this stage is not reasonable.

The mass estimates for Skylon supposedly have AIAA-standard mass growth margins built in, and that doesn't include the payload margin of slightly less than a tonne.  Take out the margins, and you might have roughly a 23-tonne payload vs. about 46 tonnes of dry mass (not including ~5 tonnes of fluids at MECO), give or take a couple of tonnes (I'm not privy to the exact values carried on each subsystem, so I assumed a global 15%).  That's about a 2:1 structure-to-payload ratio, contrasted with the usual 8:1 or worse for all-rocket hydrolox SSTO.  They're still working on firming up their mass estimates, but then AFAIK they haven't frozen the engine parameters yet...

REL has done a reasonably large number of tests and has matured some of the risk but it still introduces a lot of technology that has never been tested before. Examples would include a multi-mode engine which makes the SSME look like simplicity incarnate, a completely new aeroshell construction method, and an aerodynamic shape that has to balance subsonic performance, supersonic to low hypersonic flight in dense atmosphere, and upper atmosphere hypersonic reentry. Meanwhile, much more "traditional" SSTO proposals can get similar or higher margins and less payload uncertainty with way fewer new technologies and a more reasonable development budget... and of course for any TSTO RLV the margins go through the roof and much of the risk can be retired by turning an existing expendable launcher into an RLV.
« Last Edit: 04/21/2016 12:15 am by Nilof »
For a variable Isp spacecraft running at constant power and constant acceleration, the mass ratio is linear in delta-v.   Δv = ve0(MR-1). Or equivalently: Δv = vef PMF. Also, this is energy-optimal for a fixed delta-v and mass ratio.

Offline 93143

  • Senior Member
  • *****
  • Posts: 3054
  • Liked: 312
  • Likes Given: 1
a multi-mode engine which makes the SSME look like simplicity incarnate

Which is what they're focusing on up front because it's the key to the whole thing.  It's also not what people are typically skeptical about; note that the USAF has concluded that it will work.

As for the complexity, there's a difference between complexity in a schematic and real-world difficulty in getting something to work reliably.  The RS-25 is a significantly higher-pressure engine than SABRE, and has a lot of issues with highly stressed components in adverse thermal and chemical environments that SABRE is able to sidestep due to its unique cycle.  SABRE is more innovative, but it doesn't push the envelope as hard.

Quote
a completely new aeroshell construction method

Not new.  It was used in zeppelins.  And trusses are very easy to analyze.  Or are you talking about how they attach the actual shell to the truss?

Quote
an aerodynamic shape that has to balance subsonic performance, supersonic to low hypersonic flight in dense atmosphere, and upper atmosphere hypersonic reentry.

Turns out that's not all that tough if you know how to design an airplane, which they do.  There are some fiddly bits you need to watch out for to make reentry work, but the low ballistic coefficient helps enormously, and according to DLR's simulations, REL's analysis was pessimistic.

Why are you talking as though they haven't tackled these problems yet?  They were doing Mach 12 wind tunnel tests on the airframe shape over a decade ago.

Quote
Meanwhile, much more "traditional" SSTO proposals can get similar or higher margins and less payload uncertainty with way fewer new technologies and a more reasonable development budget...

Analysis or citation please.  This statement seems preposterous on its face.

Also keep in mind that Skylon's development budget covers everything from the £4B engine programme through roughly 400 test flights of two production prototypes (as distinct from development prototypes).  It's more like an airliner development than a rocket development.
« Last Edit: 04/21/2016 04:56 am by 93143 »

Online HMXHMX

  • Full Member
  • ****
  • Posts: 1710
  • Liked: 2215
  • Likes Given: 662


Meanwhile, much more "traditional" SSTO proposals can get similar or higher margins and less payload uncertainty with way fewer new technologies and a more reasonable development budget...

Analysis or citation please.  This statement is preposterous on its face.

Actually, if a Skylon vehicle used dense propellants, conventional high t/w rocket engines and was launched vertically, it would place somewhere between 2 and 4 times more payload into orbit for its empty weight (obviously swapping out the SABRE engines for rocket engines).  I've done this calculation for several different iterations of the published Skylon empty mass, so don't wish to bother to hunt up the link here on NASAspaceflight at the moment.  But using published info on Skylon anyone can repeat the calculation.

The calculation illustrates Skylon's real technical risk (beyond the engines working as planned).  No surprise, but its mass fraction is quite high, higher than I'd be comfortable proposing for a LOX-hydrocarbon vehicle, even VTOL.  It just looks low due to the larger amount of LH2 it employs.  But PMF calculations always must be normalized to propellant density, not weight, since that is the proper metric.

Offline 93143

  • Senior Member
  • *****
  • Posts: 3054
  • Liked: 312
  • Likes Given: 1
Okay, this is going to be handwavy:

Using Skylon C1 as a baseline, that's 1105 tonnes of propellant, vs. about 217 in the base vehicle.  To simplify the analysis, I will assume the engines are still mounted on the wingtips - ie: the vehicle still has wings, and can use them to land on the existing undercarriage (which could possibly be a couple of tonnes lighter since it isn't used for takeoff).  The tankage fraction will be better; probably not five times better, but let's assume that anyway.  The truss will, however, have to take five times the load as a first approximation, so it might weigh 20 tonnes instead of 4.  The thrust required to take off vertically should be perhaps 1400 tonnes or so, meaning a set of kerolox engines with a T/W of 180 will weigh close to 8 tonnes, down from just under 11 tonnes, and it may be possible to save a couple of tonnes on the nacelles too.  The wings don't have to lift the loaded weight of the vehicle, but they do have to take 1400 tonnes of thrust instead of 300, which could increase their mass by up to five times, to roughly 24 tonnes from a little over 5 (I don't actually expect an increase that large, since the load is mostly along the long axis, but there will probably be some).

A Merlin 1D might give an average Isp of about 302 seconds; with a total delta-V of 8.8 km/s, 1105 tonnes of propellant can put 65 tonnes in LEO.  Skylon C1 less payload weighs about 45 tonnes at MECO.  Accounting for the changes enumerated above results in an increase of between 9 and 32 tonnes, resulting in a payload between 11 and -12 tonnes in LEO.  Increasing the delta-V to 9 km/s subtracts a little over 9 tonnes from the payload, leaving it marginal in the best case.

Using RD-180 instead of Merlin might average about 329 seconds, but the engines would weigh an extra 10 tonnes or so.  With 77 tonnes in LEO, the payload is now between 13 and -10 tonnes.  Going to 9 km/s takes ~5 tonnes off that, for a payload between 8 and -15 tonnes.

Of course if you put the engines on the bottom without changing anything else, it would fold up like an accordion before you so much as lit them, so it's not trivial to remove the wings from the analysis...

I'm sure I've missed something.  1% is pretty darn good for a kerolox SSTO...  In any case I trust my point is clear...
« Last Edit: 04/21/2016 07:20 am by 93143 »

Offline Elmar Moelzer

  • Senior Member
  • *****
  • Posts: 3661
  • Liked: 849
  • Likes Given: 1062
The nuclear DC-X design research suggested that LOX-afterburning (technically TAN but I prefer to call it afterburning rather than augmentation because you're burning the oxygen and hydrogen together to increase exhaust energy, rather than just increasing mass flow) roughly triples the thrust while losing about 40% of your exhaust velocity. So for a standard solid-core NTR with 1000 s of specific impulse and a T/W ratio of around 3, you're looking at a launch T/W of 9 at 600 s of specific impulse.
Timberwind would have had a T/W of 30. Just throwing that in here.

Unfortunately, while a solid-core NTR doesn't emit fissile material, it does irradiate the H2 propellant stream enough to make it mildly radioactive.
From all I know, it is very hard to make hydrogen significantly radioactive. This is one reason why it is so often chosen as a "fuel" for NTRs. I mean the worst you can get from it is Tritium and for that you have to create Deuterium first. The neutron absorbtion cross section of protium and deuterium is extremely small.
From all I know, the only radioactive materials in the exhaust are usually small particles of the reactor fuel and those should not happen if the reactor works perfectly as designed. But then I am willing to learn new information, if you have any references.

Offline john smith 19

  • Senior Member
  • *****
  • Posts: 10351
  • Everyplaceelse
  • Liked: 2431
  • Likes Given: 13606
REL has done a reasonably large number of tests and has matured some of the risk but it still introduces a lot of technology that has never been tested before.
At full scale.
Quote
Examples would include a multi-mode engine which makes the SSME look like simplicity incarnate,
The J58 and it's nacelle (the two are deeply intwined) on the SR71 had something like 12 configurations.
Concorde had 13 analog and digital processors for each of its engines to run the inlet and exhaust configurations from takeoff to landing.

The SSME actually operates on 2 separate cycles (expander and SC) during startup. BTW there is no CAD model for the SSME. Parts of it were modeled piecemeal when problems developed. The one that now exists was built after the Shuttle programme ended.

The apparent complexity of SABRE allows REL to de-couple elements from each other in a way the SSME could not.

Building a turbine that can generate major torque from a stream of hot Helium is much easier than building one that has to resist superheated steam and hot H2.   
Quote
a completely new aeroshell construction method,
Riveting is still very much a standard method of construction and repair in the aircraft industry.
Quote
and an aerodynamic shape that has to balance subsonic performance, supersonic to low hypersonic flight in dense atmosphere, and upper atmosphere hypersonic reentry.
They started by recognizing the extreme CoG and CoL shifts that take place from launch to landing.
You seem to think this shape is untested but it's a derivative of one tested in Britain in the mid 1950s. REL have done further wind tunnel testing since then. I'd fully expect they've also used CFD to test it throughout it's flight envelope.
Quote
Meanwhile, much more "traditional" SSTO proposals can get similar or higher margins and less payload uncertainty with way fewer new technologies
And by "uncertain" you mean "smaller" ? That's a statement I can believe.

However  all  VTOL have to fit the entire structure and payload (including the engines) into a much smaller part of the GTOW.  A VTOL SSTO that delivers the same payload fraction as a TSTO ELV would push the structures and engines even harder.
Quote
and a more reasonable development budget...
AFAIK no one is currently funding a VTOL SSTO anywhere in the world. If you know of someone who has secured funding I'd be most interested to hear of them.
Quote
and of course for any TSTO RLV the margins go through the roof
Unless you use a strict Biamese or Triamese concept you need double the development budget. I'd like to some cost estimates for such a vehicle that's been put through one of the industry standard cost methodologies and see what their numbers are.
Quote
and much of the risk can be retired by turning an existing expendable launcher into an RLV.
I look forward to the first flight of SX semi reusable F9 with much interest, especially in the price they will charge for it.
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Online HMXHMX

  • Full Member
  • ****
  • Posts: 1710
  • Liked: 2215
  • Likes Given: 662
Using Skylon C1 as a baseline, that's 1105 tonnes of propellant, vs. about 217 in the base vehicle.  To simplify the analysis, I will assume the engines are still mounted on the wingtips - ie: the vehicle still has wings.  The tankage fraction will be better; probably not five times better, but let's assume that anyway.  The truss will, however, have to take five times the load as a first approximation, so it might weigh 20 tonnes instead of 4.  The thrust required to take off vertically should be perhaps 1400 tonnes or so, meaning a set of kerolox engines with a T/W of 180 will weigh close to 8 tonnes, down from just under 11 tonnes, and it may be possible to save a couple of tonnes on the nacelles too.  The wings don't have to lift the loaded weight of the vehicle, but they do have to take 1400 tonnes of thrust instead of 300, which could increase their mass by up to five times, to roughly 24 tonnes from a little over 5 (I don't actually expect an increase that large, since the load is mostly along the long axis, but there will probably be some).

A Merlin 1D might give an average Isp of about 302 seconds; with a total delta-V of 8.8 km/s, 1105 tonnes of propellant can put 65 tonnes in LEO.  Skylon C1 less payload weighs about 45 tonnes at MECO.  Accounting for the changes enumerated above results in an increase of between 11 and 32 tonnes, resulting in a payload between 9 and -12 tonnes in LEO.  Increasing the delta-V to 9 km/s subtracts a little over 9 tonnes from the payload, leaving it negative in the best case.

Using RD-180 instead of Merlin might average about 329 seconds, but the engines would weigh an extra 10 tonnes or so.  With 77 tonnes in LEO, the payload is now between 11 and -10 tonnes.  Going to 9 km/s takes ~5 tonnes off that, for a payload between 6 and -15 tonnes.

Of course if you put the engines on the bottom without changing anything else, it would fold up like an accordion before you so much as lit them, so it's not trivial to remove the wings from the analysis...

I'm sure I've missed something.  1% is pretty darn good for a kerolox SSTO...

Just a few quick notes, as I don't have time for a protracted discussion: the tank weight won't change at all.  Volume and ullage pressure – not propellant density – determines the tank mass.  And I'd use the performance and t/w of an NK-33 as a reference engine.  A VTHL LOX-Kerosene "Skylon conversion" wouldn't be any significant mass increase at burnout over a SABRE Skylon, since wing mass, rolling gear mass, and powerplant-related mass would all decline.

Offline 93143

  • Senior Member
  • *****
  • Posts: 3054
  • Liked: 312
  • Likes Given: 1
You figure a wing that can take 1400 tonnes of thrust and 200-400 tonnes of lift+gimbaled thrust (depending on neutral thrust angle) would be lighter than one that can take 300 tonnes of thrust and 330 tonnes of lift+gimbal?  Okay...  I haven't done any real analysis; as I said I'm handwaving.

NK-33 results in an engine mass a couple of tonnes heavier than the Merlin, almost identical to that of the SABRE.  Nacelle might be lighter, but then again 1400 tonnes is an awful lot of engine to cram in there while keeping the design aerodynamically nice...  And it still only gets 71 tonnes to LEO (average Isp 320 seconds using the same methodology), so it's only about four tonnes ahead of the Merlin once the extra engine mass is accounted for.

Rolling gear, sure.  I actually added that before I saw your post.

I already assumed the tank mass doesn't change, but I'm not so sure it's true; Skylon only uses a 1 bar gage pressure, and the pressure at the bottom of the LOX tanks is going to be higher than that even before liftoff.

And anyway that truss still adds a lot of mass, unless you can explain to me how a structure designed to support a little over 200 tonnes of propellant (mostly concentrated near the middle, I'll note; there's less than 70 tonnes of LH2) can manage 1100 tonnes without modification.
« Last Edit: 04/21/2016 08:01 am by 93143 »

Offline R7

  • Propulsophile
  • Senior Member
  • *****
  • Posts: 2725
    • Don't worry.. we can still be fans of OSC and SNC
  • Liked: 992
  • Likes Given: 668
You figure a wing that can take 1400 tonnes of thrust and 70 tonnes of lift would be lighter than one that can take 300 tonnes of thrust and 275 tonnes of lift?  Okay...  I haven't done any real analysis; as I said I'm handwaving.

There's no need to put NK-33s or equivalent to the wing tips. Put them in their natural place at the rear end of the vehicle and enjoy the relaxed design requirement for the wings to cope with landing lift forces only.
AD·ASTRA·ASTRORVM·GRATIA

Offline 93143

  • Senior Member
  • *****
  • Posts: 3054
  • Liked: 312
  • Likes Given: 1
And the massively increased strength requirement for the truss now that the entire ~1200-tonne vehicle is in compression.

BTW I changed my post while you were writing.  Doesn't affect your point, but...

Offline R7

  • Propulsophile
  • Senior Member
  • *****
  • Posts: 2725
    • Don't worry.. we can still be fans of OSC and SNC
  • Liked: 992
  • Likes Given: 668
IAmNotAStructuralEngineer but... pushing couple pressurized cylinders on top of each other from behind straight up sounds a lot more easier structurally than lifting two end-to-end connected horizontal cylinders from the middle while ensuring the whole thing doesn't sag and snap, even if latter cylinders are considerably lighter. But in the end I just wave hands too :)

Personally not the biggest fan of Skylon anyway. Looks cool but too non-KISS for me.
AD·ASTRA·ASTRORVM·GRATIA

Offline 93143

  • Senior Member
  • *****
  • Posts: 3054
  • Liked: 312
  • Likes Given: 1
pushing couple pressurized cylinders on top of each other from behind straight up sounds a lot more easier structurally than lifting two end-to-end connected horizontal cylinders from the middle while ensuring the whole thing doesn't sag and snap, even if latter cylinders are considerably lighter.

It isn't.  In the second case you're only pushing half the length (and mass) and the rest just hangs down from the support.  Max load is halved, as is average load.  And anyway compression is way gnarlier than tension because you can get buckling, and length makes it worse.
« Last Edit: 04/21/2016 08:13 am by 93143 »

Offline R7

  • Propulsophile
  • Senior Member
  • *****
  • Posts: 2725
    • Don't worry.. we can still be fans of OSC and SNC
  • Liked: 992
  • Likes Given: 668
1. It isn't.  In the second case you're only pushing half the length (and mass) and the rest just hangs down from the support.  Max load is halved.
2And anyway compression is way gnarlier than tension because you can get buckling, and length makes it worse.

1. Err, the read end is not going to follow you just by asking nicely. So you push half and pull the other half.
2. Remember that you are dealing with large pressurized tanks. Longitudinal tension on the tank wall is huge, in regular LVs effectively always greater than compressive loads.
AD·ASTRA·ASTRORVM·GRATIA

Offline 93143

  • Senior Member
  • *****
  • Posts: 3054
  • Liked: 312
  • Likes Given: 1
Err, the read end is not going to follow you just by asking nicely. So you push half and pull the other half.

Yes, and I'm not talking about the force exerted by the central lifting structure; I'm talking about the load in the structures it's lifting.  The peak load in the top half is right above the support - it's compressive and equal to the mass of the upper half times the acceleration induced by the supporting force (including g, of course, if applicable).  The peak load in the bottom half is right below the support - it's tensile and equal to the mass of the lower half times the acceleration.

In the fully-compressive case, the peak load is the mass of the entire vehicle times the acceleration, and it happens right at the bottom.

Quote
Remember that you are dealing with large pressurized tanks. Longitudinal tension on the tank wall is huge, in regular LVs effectively always greater than compressive loads.

This is not a regular LV.  What I'm dealing with is a spaceframe truss.
« Last Edit: 04/21/2016 08:52 am by 93143 »

Tags:
 

Advertisement NovaTech
Advertisement Northrop Grumman
Advertisement
Advertisement Margaritaville Beach Resort South Padre Island
Advertisement Brady Kenniston
Advertisement NextSpaceflight
Advertisement Nathan Barker Photography
0