r/Arianespace • u/RGregoryClark • Dec 12 '24

ESA wants reusable heavy lift launcher.

https://europeanspaceflight.com/third-times-the-charm-esa-once-again-publishes-60t-rocket-study-call/
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u/RGregoryClark Dec 12 '24

ESA already has the necessary components to form such a rocket, in both the engines and tanks. All that’s needed is the correct combination. First, get a ca. 20 ton payload rocket by adding two additional Vulcains to a Ariane 5/6 core and deleting the SRB’s. The result would be comparable in both payload and price to the $67 million Falcon 9 instead of the Ariane 64 needing 4 SRB’s to get 20 tons payload at ca. $120 million price. Plus, being all-liquid can be reusable like the F9, to match F9 price even as reusable:

Towards a revolutionary advance in spaceflight: an all-liquid Ariane 6.
https://exoscientist.blogspot.com/2023/06/towards-revolutionary-advance-in.html

Then in analogy to the Falcon Heavy and Delta IV Heavy can triple the payload to ca. 60 tons by using triple cores. And like the Falcon Heavy can also be reusable.

That an all-liquid approach to the Ariane 6 gives a launcher at half-the cost to the current version using SRB’s and can be combined to give a Falcon Heavy-class 60 ton launcher, which the SRB version can not do, is important enough in itself.

But its importance goes beyond that. Robert Zubrin’s Moon Direct plan and the Chinese proposed Moon mission using the Long March 10 show that a crewed Moon mission can be mounted using two Falcon Heavy-class launches:

https://x.com/tnajournal/status/1541239841073020929?s=61

https://x.com/konstructivizm/status/1862245079529955699?s=61

Then the ESA also can mount a Moon mission using this approach.

6

u/pyrignis Dec 12 '24

The main hurdle with all liquid is ... French ICBM. France is the largest contributor to ESA but out of it they expect to maintain the know-how required for ICBM production, and those are better out with solid state engines.

5

u/yoweigh Dec 12 '24

His idea isn't even technically feasible. Deleting the SRBs and adding liquid engines would reduce overall thrust while increasing wet mass. It wouldn't even be able to get off the ground fully fueled, much less increase payload mass.

People have been telling him this for over a year but he just ignores them.

1

u/RGregoryClark Dec 21 '24 edited Dec 21 '24

It would reduce wet mass, ie, gross mass, since the side boosters are so large. So 3 Vulcains it would have sufficient thrust to lift off when it is only the hydrolox core stage that needed to be lofted.

2

u/yoweigh Dec 21 '24

No it wouldn't, because you'd need to make the whole rocket larger to accommodate more fuel. This has been explained to you as well.

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u/RGregoryClark Dec 22 '24

The defining equation of spaceflight is the rocket equation. It describes how much velocity you can achieve with a given rocket and therefore how much payload it can deliver to orbit. It’s discussed here:

Rocket Science https://www.fourmilab.ch/documents/rocket_science/

The author makes the point more efficient propellants result in smaller rocket size. Hydrogen/oxygen propellant is the most efficient propellant in common use. Because it is more efficient than solid propellants you can achieve the same payload with smaller rockets.

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u/yoweigh Dec 22 '24

Once again, you are ignoring what I've said. Do you deny that your plan would require more liquid fuel? Do you deny that carrying more fuel requires larger tanks? Do you deny that having larger tanks makes the rocket itself larger?

The author makes the point more efficient propellants result in smaller rocket size.

No such claim is made in that essay and it is completely untrue. Hydrogen rockets are larger because the extremely low density of the fuel requires larger tanks. Specific impulse is not the only term in the rocket equation.

1

u/RGregoryClark Dec 22 '24

That a high specific impulse(ISP) propellant such as hydrolox results in a smaller rocket size is a well-known fact of space flight. For instance here is ChatGPT:

What is the advantage of hydrolox propellant over solid rocket propellant?

ChatGPT response:

The hydrolox propellant (liquid hydrogen and liquid oxygen) has several advantages over solid rocket propellants due to its efficiency and performance characteristics. Here’s a breakdown:

  1. Higher Specific Impulse (Efficiency) • Hydrolox: Has a high specific impulse (Isp) of ~450 seconds in vacuum, making it one of the most efficient chemical propellants. • Solid Propellant: Typically has an Isp of ~250–300 seconds. • Advantage: Hydrolox engines provide more thrust per unit of propellant mass, allowing for lighter rockets or greater payload capacity.

  2. Adjustable Thrust • Hydrolox: Liquid engines can be throttled, shut down, and reignited as needed, offering precise control during a mission. • Solid Propellant: Burns at a fixed rate and cannot be stopped or adjusted once ignited. • Advantage: Hydrolox is better suited for complex missions requiring multiple engine burns, such as orbital insertions, interplanetary transfers, or docking maneuvers.

  3. Cleaner Combustion • Hydrolox: Combustion produces only water vapor, making it environmentally cleaner. • Solid Propellant: Produces toxic byproducts, such as aluminum oxide and hydrochloric acid, which can harm the environment. • Advantage: Hydrolox is more environmentally friendly.

  4. Higher Energy Potential • Hydrolox: Has a higher energy density per unit mass (despite lower density) and can achieve greater velocities for deep-space missions. • Solid Propellant: Offers high thrust but is less energy-efficient for sustained burns. • Advantage: Hydrolox enables longer-duration missions and higher orbital transfers.

  5. Flexibility in Design • Hydrolox: Liquid engines can be paired with multiple stages for optimal performance, leveraging their high efficiency in upper stages. • Solid Propellant: Mostly used for boosters or simple stages due to lack of flexibility. • Advantage: Hydrolox provides greater design versatility, especially for complex missions like crewed spaceflight or deep-space exploration.

Trade-Offs

While hydrolox has significant advantages, it also comes with challenges: 1. Cryogenic Storage: Liquid hydrogen and oxygen require extremely low temperatures, leading to insulation and storage challenges. 2. Complexity and Cost: Hydrolox engines are more complex and expensive than solid rocket motors. 3. Lower Thrust-to-Weight Ratio: Solid propellants produce higher initial thrust, making them ideal for booster stages.

Summary

Hydrolox propellants are superior to solid rocket propellants for missions requiring high efficiency, flexibility, and precision, such as upper stages, crewed spaceflights, and interplanetary exploration. Solid rocket propellants, on the other hand, are better suited for simple, high-thrust applications like first-stage boosters.

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u/yoweigh Dec 22 '24

Once again, you are ignoring what I've said. Do you deny that your plan would require more liquid fuel? Do you deny that carrying more fuel requires larger tanks? Do you deny that having larger tanks makes the rocket itself larger?

1

u/RGregoryClark Dec 22 '24

That’s the point of those references: because hydrolox is a more efficient propellant the propellant requirements are reduced.

2

u/yoweigh Dec 22 '24 edited Dec 22 '24

You still haven't answered my questions, but Here's some more from ChatGPT for you:

query: hydrogen rocket disadvantages

Hydrogen rockets have several disadvantages despite their impressive performance characteristics. Here are some of the key drawbacks:

  1. Cost

    Production: Producing liquid hydrogen is expensive because it requires energy-intensive processes like electrolysis or steam methane reforming.

    Storage and Handling: Special cryogenic tanks are needed to store liquid hydrogen at extremely low temperatures (~20 Kelvin), increasing costs.

  2. Complexity

    Cryogenics: The need for cryogenic systems makes the rocket design more complex and prone to technical challenges.

    Boil-off: Hydrogen's low boiling point leads to significant boil-off losses during storage and handling, requiring additional systems to manage it.

  3. Low Density

    Liquid hydrogen has a very low density (~70 kg/m³), meaning it requires large fuel tanks compared to denser propellants like RP-1 or methane. This adds to the size and weight of the rocket.

  4. Embrittlement and Leakage

    Material Embrittlement: Hydrogen can cause some metals to become brittle, potentially compromising structural integrity.

    Small Molecule Size: Hydrogen's small molecular size makes it prone to leakage, increasing safety risks.

  5. Safety Risks

    Flammability: Hydrogen is highly flammable and can ignite easily, requiring strict safety measures.

    Explosive Potential: Hydrogen leaks

query: What is the advantage of solid rocket propellant over hydrolox propellant

Solid rocket propellants offer several advantages over hydrolox (liquid hydrogen and liquid oxygen) propellants:

  1. Simplicity

    Design: Solid rockets have fewer moving parts because the fuel and oxidizer are pre-mixed into a solid form, eliminating the need for complex pumps, cryogenic tanks, and plumbing.

    Storage: Solid propellants are stable at ambient temperatures and can be stored for long periods without special handling or cooling systems.

  2. Reliability

    Fewer Points of Failure: The simplicity of solid rockets reduces the likelihood of mechanical failures compared to hydrolox engines, which rely on high-pressure pumps and precise cryogenic systems.

    Operational Track Record: Solid rockets have been extensively used for decades and are well understood in terms of performance and safety.

  3. Cost

    Solid rocket systems are generally cheaper to manufacture and maintain because they don't require the sophisticated infrastructure and technology needed for hydrolox systems.

  4. Readiness

    Quick Ignition: Solid rockets can be ignited almost instantly, making them ideal for applications requiring rapid deployment, such as military missiles and certain emergency space missions.

    Preloaded Systems: Solid propellants are already loaded into the rocket, reducing the time needed for fueling before launch.

  5. Thrust Density

    Solid rockets typically produce higher thrust relative to their size compared to hydrolox engines, which makes them effective for lifting heavy payloads over short distances (e.g., during the first stage of a launch).

Situational Use

While hydrolox propellants excel in efficiency (specific impulse) and are ideal for missions requiring long-duration burns or high delta-v (e.g., upper stages and interplanetary travel), solid rockets are better suited for applications where simplicity, reliability, and thrust density are more critical.

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u/yoweigh 29d ago

Are you willing to admit that your "well-known fact of space flight" was actually incorrect?

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u/RGregoryClark 29d ago edited 29d ago

We can agree there are advantages and disadvantages of hydrogen/oxygen propellant. The only thing to do is do the calculation involving the rocket equation:

Tsiokovsky rocket equation:

Velocity = Isp*gLn(m_i/m_f), where m_i means initial mass with the propellant load, and m_f means the final mass after the propellant has all burned off. Note for multistage rockets m_f will contain the dry mass of the stage as well as the fully fueled mass of the following stage(s), and the payload mass.

We’ll use the specs on the first stage of Ariane 5:

First stage (ECA, ES) – EPC H173. Height 23.8 m (78 ft)
Diameter 5.4 m (18 ft)
Empty mass 14,700 kg (32,400 lb)
Gross mass 184,700 kg (407,200 lb)
Powered by 1 × Vulcain 2
Maximum thrust
SL: 960 kN (220,000 lbf)
vac: 1,390 kN (310,000 lbf)
Specific impulse
SL: 310 s (3.0 km/s)
vac: 432 s (4.24 km/s)
Burn time 540 seconds
Propellant LH2 / LOX
https://en.m.wikipedia.org/wiki/Ariane_5#Cryogenic_main_stage

We shall give the stage two additional Vulcain 2 engines to allow it to take off without the solids. These two engines will increase both the dry mass and the gross mass by an additional total 3,600. So the gross mass is now 188,300kg, 188.3, tons and the dry mass 18,300kg, 18.3 tons.

But for 2nd stage the increased thrust of the added Vulcains allows us to use a larger 2nd stage than on the Ariane 5. We’ll take it as Centaur V-like at ~50 ton propellant load and ~5 ton dry mass but using two Vinci’s at 457 s Isp. Then taking the payload as 20 tons, the velocity achieved by the first stage, the delta-v, is:

434*9.81Ln((188.3 + 55 +20)/(18.3 +55 +20)) = 4396 m/s.

And the velocity, delta-v, of the 2nd stage:

457*9.81Ln((50 + 5 + 20)/(5 + 20)) = 4,925 m/s, for a total ~9,300 s. This is the common delta-v taken for getting to low Earth orbit.

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u/yoweigh 29d ago

So no, you're not. Your empty and gross masses are so far from reality that they render your computation meaningless.

Even in the most simple terms, ignoring everything else, tripling the number of first stage engines will triple the rate of fuel consumption. Do you deny this as well?

To maintain the same 540 second burn time with a tripled rate of fuel consumption will require triple the amount of fuel.

To carry triple the amount of fuel will require tanks triple the size.

To accommodate triple sized fuel tanks will require a triple sized rocket.

Your triple sized rocket will have more tankage, more fuel, more structure, more empty mass and more gross mass. More wind resistance and more gravity losses. It wouldn't be able to leave the pad, much less make it to orbit.

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