New start-up aiming for a SSTO.

[Exoscientist]

  Bob Clark

[AckSed]

10-kg payload to SSO, 13 kg to LEO, launch anywhere... It's the Mockingbird!

Okay, so a while back I saw a concept for a very small SSTO designed by Lawrence Livermore: https://up-ship.com/blog/?p=11056

It was initially designed to be a target drone, but there was enough delta-V that the drone could reach orbit and lob a 10kg payload into LEO, hence its nickname: "Bricklifter".

It goes without saying that the margins and potential for weight-gain were and are severe, especially with reuse added on. However, these days a 10kg cubesat can be pretty damn capable. For reference, a 1U cubesat can mass up to 2kg. A 6U cubesat should be possible, and the ESA is very interested in cubesats.

From the images, it looks like they're using something cryogenic (I see some cylinders in the integrated test that look like oxygen), another fuel (orange flame, so some form of kerosene) and stainless steel balloon tanks. Mockingbird was aluminium, but smaller.

I think that shroud on the bottom in the proposed launch diagram is part of the heatshield. Whether they're expecting the light structure to just tank it or include active cooling, I don't know.

Honestly, if they pull this off, fair play to them. They could do a good bit of business out of it.

[sevenperforce]

It's a kerolox rocket that uses an aft-first entry heat shield and parafoil recovery:

More power to them if they can make it work! Kerolox doesn't offer much margin here.

It looks like they are contemplating a cluster of 7-9 engines inside the toroidal heat shield, possibly relying on stagnation inside the heat shield to protect the engines. It's claiming 25 kN of thrust which, for an assumed 1.3 T/W ratio, caps the liftoff mass at around 2 tonnes. This looks like a pressure-fed engine so I don't see how they get a vacuum specific impulse greater than 325 seconds or so. To achieve over 9300 m/s of Δv, it will need a dry mass no greater than 106 kg, of which 13 kg is the payload.

That's pretty tight.

[Exoscientist]

 A speculation on a small SSTO based on the Falcon 9 upper stage. Here’s the specifications of the F9 upper stage:

 

 The Merlin Vacuum on the 2nd stage can’t operate at sea level because of its high expansion. And the sea level Merlin would not have enough thrust to loft the stage from the ground. 
 So I’ll reduce the propellant load by 1/2. A single sea level Merlin could then launch it. I’ll estimate this half-size 2nd stage’s dry mass by first subtracting off the engine mass from the dry mass, taking half the remaining mass, then adding back on the mass of the engine. The reason is you still need the full engine size and mass to lift off, not a half-size engine.
 This can only be an estimate of the dry mass since not all components will reduce in size proportionally.

 But with this estimated dry mass the results are as below:

 

 

 So about 670 kg to orbit as an expendable. For an operational SSTO you really want to use altitude compensation. I estimate using it you could raise the payload to ca. 1,600 kg or possibly higher.

Another nice thing about this is you could get 10 of the these small SSTO’s from one Falcon 9, since there are 10 Merlins on the Falcon 9, and the total propellant size on the Falcon 9 of 500 tons amounts to 10 of the small size SSTO’s.

So you could buy a single reused Falcon 9 at $40 million, and break it down to 10 of the small SSTO’s at $4 million each.

  Bob Clark

Edited July 3 by Exoscientist

[AckSed]

There might be a bit more margin to squeeze out. I'll re-post the sci.space.tech USENET post on the differences between hydrolox and RP-1/H2O2 SSTO: https://yarchive.net/space/rocket/fuels/hydrogen_deltav.html

The key takeaway is that, given the same mount of thrust, denser fuels with lower vacuum ISP burn out faster. The vehicle thus lightens faster and accelerates more quickly, suffering less gravity loss even if you have to limit the G-forces. As such, it takes the RP-1/H2O2 SSTO slightly less delta-V to reach their reference ISS LEO orbit - 8855 m/s. And the stated vacuum ISP was 320 seconds.

This is comparing pears to apples, but the key point is a kerelox SSTO doesn't need as much delta-V to reach orbit as you think.

[Exoscientist]

On 7/2/2024 at 12:50 PM, AckSed said:

10-kg payload to SSO, 13 kg to LEO, launch anywhere... It's the Mockingbird!

Okay, so a while back I saw a concept for a very small SSTO designed by Lawrence Livermore: https://up-ship.com/blog/?p=11056

It was initially designed to be a target drone, but there was enough delta-V that the drone could reach orbit and lob a 10kg payload into LEO, hence its nickname: "Bricklifter".

It goes without saying that the margins and potential for weight-gain were and are severe, especially with reuse added on. However, these days a 10kg cubesat can be pretty damn capable. For reference, a 1U cubesat can mass up to 2kg. A 6U cubesat should be possible, and the ESA is very interested in cubesats.

From the images, it looks like they're using something cryogenic (I see some cylinders in the integrated test that look like oxygen), another fuel (orange flame, so some form of kerosene) and stainless steel balloon tanks. Mockingbird was aluminium, but smaller.

I think that shroud on the bottom in the proposed launch diagram is part of the heatshield. Whether they're expecting the light structure to just tank it or include active cooling, I don't know.

Honestly, if they pull this off, fair play to them. They could do a good bit of business out of it.

Thanks for the Mockingbird SSTO link.  This though did use pump-fed engines. If Sidereus plans pressure-fed then their plan is much more problematical because of heavy tank weight. 

 The pumps Mockingbird used though were piston pumps, since turbo pumps are harder to get to work at small size. 

 The XCOR company showed piston-pump rocket engines are doable since they built them and actually did test flights with them on rocket propelled airplanes:

XCOR Aerospace.
https://kmhv.wordpress.com/

 XCOR went bankrupt but their assets were purchased by a space-education company:

Bankrupt Spaceflight Company's Space Plane Assets to Help Young Minds Soar
News
By Douglas Messier published April 20, 2018
https://www.space.com/40352-xcor-aerospace-lynx-space-plan-stem-education.html

  Then Sidereous could purchase these engines ready made. 

  Bob Clark

Edited July 4 by Exoscientist

[kerbiloid]

On 7/2/2024 at 7:50 PM, AckSed said:

10-kg payload to SSO, 13 kg to LEO, launch anywhere...

For then-planned "Laika" dawg it was set a limit of 6..7 kg.

So, enough for a dawgnaut with a small air balloon.

Or with a small inflatable heatshield, depending on what it prefers: to breathe or to land.

Edited July 4 by kerbiloid

[magnemoe]

23 hours ago, sevenperforce said:

It's a kerolox rocket that uses an aft-first entry heat shield and parafoil recovery:

More power to them if they can make it work! Kerolox doesn't offer much margin here.

It looks like they are contemplating a cluster of 7-9 engines inside the toroidal heat shield, possibly relying on stagnation inside the heat shield to protect the engines. It's claiming 25 kN of thrust which, for an assumed 1.3 T/W ratio, caps the liftoff mass at around 2 tonnes. This looks like a pressure-fed engine so I don't see how they get a vacuum specific impulse greater than 325 seconds or so. To achieve over 9300 m/s of Δv, it will need a dry mass no greater than 106 kg, of which 13 kg is the payload.

That's pretty tight.

¨Wait it only weight 2 ton? this give the option to air lift, my guess is that its to wide to drop from an fighter jet who let it drop with close to supersonic, climbing and high up. 
Is this not pretty close to the smallest orbital rocket, it was dropped from an fighter jet, two stages and obviously disposable. 
For tiny rocket air resistance is much worse as diameter is huge compared to weight so an slimmer design might be better but then landing is harder. 

[Exoscientist]

On 7/4/2024 at 11:59 AM, magnemoe said:

¨Wait it only weight 2 ton? this give the option to air lift, my guess is that its to wide to drop from an fighter jet who let it drop with close to supersonic, climbing and high up. 
Is this not pretty close to the smallest orbital rocket, it was dropped from an fighter jet, two stages and obviously disposable. 
For tiny rocket air resistance is much worse as diameter is huge compared to weight so an slimmer design might be better but then landing is harder. 

 

 Good point.  Assuming it is similar to the Mockingbird design that @AckSed linked, https://up-ship.com/blog/?p=11056, then it would have a dry mass in the range of 75 kg and a gross mass in the range of 1,500 kg. 

 DARPA had wanted an air-launched system for small payloads at $1 million launch cost for a 45 kg payload with their ALASA program:

 

The ALASA rocket, measuring 7.3 meters long, would be attached to the underbelly of a Boeing-built F-15E fighter aircraft. DARPA says taking off from a standard airport runway would allow the Defense Department to launch from almost anywhere. Credit: Boeing artist's concept.
https://spacenews.com/40023boeing-targets-66-percent-launch-cost-reduction-with-alasa/

 See discussion here:

Dave Masten's DARPA Spaceplane, page 2: an Air Launched System.
http://exoscientist.blogspot.com/2014/08/dave-mastens-darpa-spaceplane-page-2.html

 Boeing initially was investigating the possibility but pulled out of the program. I believe they pulled out not because they couldn’t get it to work, but because they couldn’t make enough money on it. (See my sig file.)

 But at the specifications of Mockingbird, and perhaps Sidereus, it could easily make the DARPA requirements. This research report investigated the advantages of air-launch:

A.4.2.1 Launch Method Analysis (Air Launch).
For a launch from a carrier aircraft, the aircraft speed will directly reduce the Δv required to attain LEO. However, the majority of the Δv benefit from an air launch results from the angle of attack of the vehicle during the release of the rocket. An ideal angle is somewhere of the order of 25° to 30°.
A study by Klijn et al. concluded that at an altitude of 15250m, a rocket launch with the carrier vehicle having a zero launch velocity at an angle of attack of 0° to the horizontal experienced a Δv benefit of approximately 600 m/s while a launch at a velocity of 340m/s at the same altitude and angle of attack resulted in a Δv benefit of approximately 900m/s. The zero launch velocity situations can be used to represent the launch from a balloon as it has no horizontal velocity.
Furthermore, by increasing the angle of attack of the carrier vehicle to 30° and launching at 340m/s, a Δv gain of approximately 1100m/s was obtained. Increasing the launch velocity to 681m/s and 1021m/s produced a
Δv gain of 1600m/s and 2000m/s respectively.
From this comparison, it can be seen that in terms of the Δv gain, an airlaunch is  superior to a ground launch. As the size of the vehicle decreases, this superiority will have a larger effect due to the increased effective drag on the vehicle.
https://web.archive.org/web/20120229141110/https://engineering.purdue.edu/AAE/Academics/Courses/aae450/2008/spring/report_archive/reportuploads/appendix/propulsion/A.4.2.1 Launch Method Analysis (Air Launch).doc

 The 681 m/s launch velocity mentioned is about Mach 2, and the 1021 m/s velocity about Mach 3. The F-15 jet fighter planned for the ALASA air-launch could get about Mach 2.5, So estimate the delta-v savings as intermediate between the 1600 m/s and 2000 m/s mentioned so about 1,800 m/s.

 Taking the required delta-v to orbit as about 9,200 m/s, the small orbital rocket would only have provide about 7,400 m/s delta-v.  For launch at high altitude take the Isp as close to the vacuum Isp of, say, the Merlin Vacuum of 348s.

 By the way, arguments that air-launch wouldn’t offer much delta-v benefit looked only at the aircraft speed but another key benefit is that launching at high altitude allows near vacuum Isp for the orbital stage, rather than sea level Isp.

 Then the Mockingbird/Sidereus rocket could get about 100 kg to orbit:

348*9.81Ln(1 + 1,425/(75 + 100)) = 7,550 m/s.

 Note the 10 times higher payload capability by using air-launch.

  Since DARPA wanted 45 kg at $1 million per launch, they might agree to pay $2 million for this much payload.

 The operational cost for the F-15 would be comparatively low cost compared to that:

USAF: The F-15EX Costs Less than the F-35 | TURDEF
The planes cost $80 million each, which is cheaper than the F-35. The USAF will receive another six units this year from Lot 1 production. The F-15EX is designed to last up to 20,000 flight hours at the cost of $29,000/hour. According to the report, this cost is a third of the operating cost of the F-35. Jan 20, 2023
https://turdef.com/article/usaf-the-f-15ex-has-a-higher-advantage-and-costs-less-than-the-f-35

 Since Sidereus was planning to offer their rocket at only $100,000 this would result in a quite high profit margin for them if DARPA paid them $2 million for the higher payload air-launched version.

   Bob Clark

Edited July 9 by Exoscientist
Added image.

[Exoscientist]

On their LinkedIn page, Sidereus said their engines are pump fed:

https://www.linkedin.com/posts/luca-principi-2801981a9_cold-flow-tests-on-the-eos-mr-5-engine-activity-7214905823716646912-0eqv

 The company co-founder gave an interview on Next Big Future about their rocket:

 

 Bob Clark

[AckSed]

On 7/9/2024 at 3:44 PM, Exoscientist said:

On their LinkedIn page, Sidereus said their engines are pump fed:

https://www.linkedin.com/posts/luca-principi-2801981a9_cold-flow-tests-on-the-eos-mr-5-engine-activity-7214905823716646912-0eqv

 The company co-founder [Mattia Barbarossa] gave an interview on Next Big Future about their rocket:

https://www.youtube.com/watch?v=Cbs1nIzj12c

 Bob Clark

Alright, this is more like the Mockingbird than I thought.

First off, its tanks are aluminium, not steel, and supposedly masses 20kg, and they're hoping to shave it down to 15kg. The oxygen tanks have withstood "more than 50 tons of force" at cryogenic temps. Assuming that's tons/m2, that's ~5 bar. If it's pump-fed, that's about right.

Second, "the entire rocket system is 10kg, more or less". Now whether that's each engine, a cluster of nozzles attached Soviet-style to a single combustion chamber or 7 separate little rocket engines in a cluster, I'm not sure.

Ambition is to make "the personal computer of space" and aims for $100,000 to orbit.

Chose SSTO because it's simpler in terms of logistics, despite being a "chimera".

Mentions the Atlas that launched Mercury as a "stage and a half". (Knows his history at least.)

Render of Eos shows 7 nozzles flat against a barrel-shaped heatshield.

Shows off RHOMBUS transport version and Star-Raker refuelling at airport (as concepts of accessible spaceflight).

Unless we make LEO as accessible as every other place on Earth, we are still not getting the interplanetary civilisation we're looking for.

Probably the smallest [orbital] rocket ever seen. It could fit inside the fairing of the Falcon 9 with copious room to spare. it's shorter and narrower than a single Raptor Vacuum engine.

Big crunchy numbers incoming:

• 1792kg gross Lift-off weight (GLOW)
• 30kg dry weight
• 41.4 to 1 mass ratio
• Jet-A1/oxygen propellant (supposedly carbon-neutral Jet-A)
• Sea-level thrust 25kN
• Sea-level ISP 260s
• Vacuum ISP 310s
• 250:1 thrust-to-weight ratio
• 4.2m tall
• 1m diameter at base
• 1 engine config with gimballing
• 12kg payload to SSO
• Reusable 10 times

(I'm taken aback by the gimballing in an engine this small and this constrained in mass and engine-bay volume. How?)

It's not as powerful as a 'proper' rocket but this, again, is a personal computer to give launch capability from a concrete stand. Launch from 1km area, deploy payload, recover.

Mentions student rocket teams, presumably in comparison to ease of launch.

10 could fit inside the F9 fairing. Could be a high-energy stage to get a small payload to Pluto.

That dark barrel-looking section in the rear of the render? That's where the payload bay and the parafoil is situated.

Mission profile:

• Max-Q 65 seconds
• MECO 420 seconds
• Payload deploy 15 minutes
• Re-entry 1 hour, 6 minutes
• Gliding to a landing 12 hours, 6 minutes
• Other missions possible: VLEO and high elliptical orbit. 1U cubesat could be launched into 185km x 12,000km orbit

Indended to fuel and go, launching with a "laptop and launch" architecture and satellite telemetry. Recovery through splashdown.

If you're intending, in the far future, to launch tens of thousands of these, you need some way to make this safe. Intent is, in case of engine shutdown or other problems, to vent the propellant and deploy parachute.

MR5 engines developed in-house. 3D-printed, "region-cooled" with Jet-A, as it's a much more accessible fuel. Each engine is in the ballpark of a 30cm water bottle in size but generates 2.5 tons thrust.

Test stand near Naples, Italy is literally a concrete pit and a few garden sheds in a field.

7 million euro in funding.

The prototype chamber is 5kg in weight. Hoping for a low-altitude flight by end of year, and sub-orbital by end of 2025. First orbital flight by 2026, more or less.

Edited July 10 by AckSed

[Exoscientist]

40 minutes ago, AckSed said:

• 1792kg gross Lift-off weight (GLOW)
• 30kg dry weight
• 41.4 to 1 mass ratio

 

 This part must be what they were just “hoping” for. If true, it would mean a mass ratio of the rocket itself aside from the payload of 60 to 1. That’s not reasonable. There is also the fact the 20 kg tank plus 10 kg engine section would already be 30 kg with no mass left over for any other systems. Plus, running the numbers, with a dry mass that low, you would get payload to LEO in the range of 50 to 60 kg, not the ~10 kg they are claiming. The Mockingbird rocket had a more realistic 20 to 1 to mass ratio, with a 75 kg dry mass.

  Bob Clark

[magnemoe]

Instinct reading SSTO, no do not do. 
So many has chased the golden unicorn who is the SSTO as I feel its better to treat an chemical rocket SSTO as an bad idea. 

Reuse is way harder, outside the shuttle and some capsules who was refurbished no upper stages has been reused. 
Currently one rocked reuse first stage  and they own the launch marked. 
 

[AckSed]

24 minutes ago, Exoscientist said:

This part must be what they were just “hoping” for. If true, it would mean a mass ratio of the rocket itself aside from the payload of 60 to 1. That’s not reasonable. There is also the fact the 20 kg tank plus 10 kg engine section would already be 30 kg with no mass left over for any other systems. Plus, running the numbers, with a dry mass that low, you would get payload to LEO in the range of 50 to 60 kg, not the ~10 kg they are claiming. The Mockingbird rocket had a more realistic 20 to 1 to mass ratio, with a 75 kg dry mass.

RE: payload being too high. I feel that's propellant for deorbit?

Well 30kg total is what they are claiming. There might also be some inexpertise with English at work here; later he says it's the engines and tanks that are 30kg. For now I will be neutral on whether it's achievable. We'll see.

More details:

Aiming for fuelling and launching within five hours.

The host points out that - extending the personal computer analogy - if you have a spacecraft of your own, you don't need an independent satellite bus like on other rideshare missions. The rocket itself has telemetry, power and so on, and utilising that increases the effective payload.

Later on, Mattias says that, massing 30kg while having an almost 1 square metre cross-section, their ballistic coefficient is extremely low, two to three orders of magnitude smaller than a garden variety reentry capsule, slowing them very quickly very high in the atmosphere, hence their entry constraints are much milder. However, it also means they cannot linger in LEO for long - six months is stated - and EOS' avionics are comparable to a cubesat, which also are sensitive to radiation. Not ruling out radiation-hardened components for specific customers.

Not using a plug nozzle like Philip Bono, but nozzle extension does much the same and serves as radiative heatshield.

Mass-optimised, not ISP-optimised.

Intend to use a multiple skip trajectory - enter very shallow angle, heat up the heatshield, come back out, radiate the heat away, dip back in. Do this a few times until you can gain no more lift, then dive in. (Probably why re-entry to landing is so long.)

Terminal velocity is just 100 km/h.

Landing intended to be a splashdown in water, as the cost of refurbishment from seawater is minute compared to more advanced recovery systems. The autonomously-steered parafoil can ensure reasonably pinpoint landings. (Not out of the question. Here's a hobbyist effort: https://hackaday.io/project/176621-r2home/details)

[AckSed]

Uncovered some more details about Mockingbird: https://yarchive.net/space/rocket/small_rocket_inefficiencies.html

Quote

Mockingbird got around the turbopump problem [of small rockets] by using reciprocating pumps, whose specific mass *improves* at smaller sizes, with a crossover relative to turbopumps somewhere in the general ballpark of 20,000 N (~4000 lbf) thrust (highly implementation and technology dependent). Mockingbird baseline engine design was regen-cooled peroxide/kerosene with a pump+chamber mass (8 chambers) of 13.3 kg and a vacuum thrust of 3.2 kN per chamber (25.6 kN total) for a T/W of 188. (Sea level thrust was 2.5 kN per chamber, 20 kN total, T/W 147)

I note the sea-level thrust and T/W ratio of Mockingbird is quite a bit less than EOS. If Siderus are forced to pack on the kilos, there may be slack to make it into a non-reusable system.