Dr. Robert Zubrin Moon Direct: How to build a moonbase in four years

SpaceNut · 2025-08-17 14:26:59

I think that the faith of using the starship and BFR has begun to lose favor seems that it might be time to rethink what can be done with a Facon 9 and Heavy combinations to build an orbital platform and path to Moon and Mars from the pieces.

We recent talked about using the first stage for a build on orbit from its fuel tanks but why not design parts with in the pieces instead.

1) mass of fuel
2) mass of vehicle
3) performance of engines
4) ISP of fuel

First stage Mass ('dry' without propellant) 22,200 kg (48,900 lb) Second stage Mass (without propellant) 4,000 kg (8,800 lb) First stage Mass ('wet' with propellant) 433,100 kg (954,800 lb) Second stage Mass (with propellant) 111,500 kg (245,800 lb)Dec 3, 2020

So we can launch an altered first stage as a habitat on top of a modified second stage.

SECOND STAGE
Height 13.8 m / 45.3 ft
Diameter 3.7 m / 12.1 ft
Empty Mass 3,900 kg / 8,598 lb

Seems more than doable.

moon direct a cost effective plan to enable human lunar exploration

what are the requirements to leverage from space x and other what can be done on orbit assembly.

To deliver program success on budget and schedule, project managers develop a common sense list of subsidiary requirements, such as:
– Minimize cost
– Minimize weight
– Minimize fuel
– Minimize risk to astronauts
– Minimize programmatic risk (e.g. getting canceled)
– Emplace infrastructure for a future series of missions
– Avoid technological dead ends (e.g. expendable vs reusable rockets)
– Avoid cost of new rocket (to avoid sticker shock, since rocket development is considered to be expensive)
– Placate constituencies (to avoid sniping)

Some of these requirements are easier to measure. Others are effectively beyond the project manager’s control. Still others give very different answers depending on the order of priority, such as the electric car example above. In a sentence, this varied order and weighting of common sense requirements is what gives rise to such different architectures in an immature product space.

SpaceNut · 2025-08-18 15:27:40

https://launchercalculator.com/

Googling for more information,
Moon Direct: A Purpose-Driven Plan to Open the Lunar Frontier [R.Zubrin]

Abstract for moon direct

Thoughts on [R.Zubrin Op-Ed] Lunar Gateway or Moon Direct? [4.17.19]

moon direct a cost effective plan to enable human lunar exploration

Table 1. Cargo Lander Mission (single stage)
Launcher Staging Orbit Propulsion Tank Length Payload Delivered
Falcon H LEO LOx/CH4 3.2 m 8.3 tons
Falcon H GTO LOx/CH4 1.05 8.3
Falcon H LEO LOx/H2 7.9 10.4
Falcon H GTO LOx/H2 2.5 9.6
New Glenn LEO LOx/CH4 1.12 6.0
New Glenn LEO ` LOx/H2 2.85 7.5
Vulcan LEO LOx/CH4 1.54 4.0
Vulcan LEO LOx/H2 3.8 5.0
SLS LEO LOx/CH4 1.9 12.0
SLS LEO LOx/H2 4.45 15.0
BFR LEO LOx/CH4 3.2 19.9
BFR TLI LOX/CH4 2.5 60.0

SpaceNut · 2025-08-18 16:19:44

A Purpose-Driven Plan to Open the Lunar Frontier

We can estimate the weight of this vehicle by considering the Apollo Lunar Module (LM). As noted, our LEV will have a similar profile:
lightweight; intended to fly only in space and around the lunar surface, meaning it would not need a thick shell and heavy heat shield to protect it during re-entry into Earth’s atmosphere; and capable of carrying some
cargo, a crew of two, and life support for up to a few days.
The Apollo LM’s dry mass (its weight with crew and cargo but without fuel) was 5.2 metric tons. However, the LM carried two rocket engines and propulsion systems—one for descending from lunar orbit onto the Moon’s surface, another for ascending back to lunar orbit. The ascent portion (the “ascent stage”), which contained the crew cabin, crew, and life support equipment, is most similar to our purposes here. Its dry mass was 2.3 tons. If we used this figure for estimating the weight of our LEV, we would also need to add the weight of the landing legs, and make various other adjustments. But given a half-century of improvements in materials and avionics science and engineering, a LEV could surely make significant improvements in the weight. We will therefore estimate 2 tons for the LEV’s dry mass, again, including crew and cargo.
In Figure 2 (see page 37), we can see the mass requirements of our 2-ton LEV. In addition to the dry mass, about 6 tons of propellant are required for each mission that uses 6.1 km/s of delta-V. So the total mass (known as the “wet mass”), including ship, cargo, and propellant, is about 8 tons. Also, the required weight of the tanks and engines—which take up part of the 2-ton dry mass—still leaves 1.3 tons for the crew, crew cabin, and other cargo.

Of course there is the refueling and other discusions that keep cost down.

SpaceNut · 2025-08-18 16:24:28

0312-Falcon-Heavy-on-Pad.jpg?w=1000&ssl=1

The Falcon Heavy can lift 60 tons to low Earth orbit (LEO). Starting from that point, a hydrogen/oxygen rocket-propelled cargo lander could deliver 12 tons of payload to the lunar surface.
We therefore proceed by sending two such landers to our planned base location. The best place for it would be at one of the poles, because there are spots at both lunar poles where sunlight is accessible all the time, as well as permanently shadowed craters nearby where water ice has accumulated. Such ice could be electrolyzed to make hydrogen-oxygen rocket propellant, to fuel both Earth-return vehicles as well as flying rocket vehicles that would provide the lunar base’s crew with exploratory access to most of the rest of the moon.
The first cargo lander carries a load of equipment, including a solar panel array, high-data-rate communications gear, a microwave power-beaming set up with a range of 100 kilometers, an electrolysis/refrigeration unit, two crew vehicles, a trailer, and a group of tele-operated robotic rovers. After landing, some of the rovers are used to set up the solar array and communications system, while others are used to scout out the landing area in detail, putting down radio beacons on the precise target locations for the landings to follow.
The second cargo lander brings out a 12-ton habitation module, loaded with food, spare spacesuits, scientific equipment, tools, and other supplies. This will serve as the astronauts’ house, laboratory, and workshop of the moon. Once it has landed, the rovers hook it up to the power supply and all systems are checked out. This done, the rovers are redeployed to do detailed photography of the base area and its surroundings. All this data is sent back to Earth, to aid mission planners and the science and engineering support teams, and ultimately forming the basis of a virtual reality program that will allow millions of members of the public to participate in the missions as well.

So to get a TSTO launch with 63,000 kg payload to LEO we need

To calculate the performance of a Two-Stage-To-Orbit (TSTO) rocket launching a payload of 63,000 kg to Low Earth Orbit (LEO), we can use the Rocket Equation (Tsiolkovsky's equation) and some basic principles of orbital mechanics. Here's the general approach:
Tsiolkovsky Rocket Equation
The rocket equation is:
$$\Delta v = v_e \cdot \ln\left(\frac{m_0}{m_f}\right)$$
Where:
$$\Delta v$$ = Change in velocity required (orbital velocity + losses)
$$v_e$$ = Effective exhaust velocity of the rocket (specific impulse $$I_{sp}$$ in seconds multiplied by $$g_0$$, the gravitational acceleration at Earth's surface, $$9.81 , \text{m/s}^2$$)
$$m_0$$ = Initial mass of the rocket (including fuel, structure, and payload)
$$m_f$$ = Final mass of the rocket (structure + payload)
Steps to Solve for TSTO
Determine Required $$\Delta v$$:
For LEO, the typical $$\Delta v$$ is approximately 9.3–10 km/s, accounting for:
Orbital velocity (~7.8 km/s)
Gravity losses (~1.5 km/s)
Drag losses (~0.2 km/s)
Stage Splitting:
TSTO rockets divide the total $$\Delta v$$ between the two stages. A common split is:
First stage: ~60–70% of $$\Delta v$$
Second stage: ~30–40% of $$\Delta v$$
Mass Ratios:
For each stage, calculate the mass ratio $$\frac{m_0}{m_f}$$ using the rocket equation.
Payload Fraction:
The payload fraction is the ratio of payload mass to the total initial mass of the rocket:
$$\text{Payload Fraction} = \frac{\text{Payload Mass}}{m_0}$$
Iterative Design:
Adjust parameters like specific impulse ($$I_{sp}$$), structural mass fraction, and fuel mass fraction to meet the payload requirement.
Example Calculation
Assume:
Payload: $$m_{\text{payload}} = 63,000 , \text{kg}$$
Specific impulse ($$I_{sp}$$):
First stage: $$300 , \text{s}$$
Second stage: $$450 , \text{s}$$
Structural mass fraction: $$0.1$$ (10% of total mass is structure)
Total $$\Delta v$$: $$9.5 , \text{km/s}$$
First Stage:
$$\Delta v_1 = 0.6 \cdot 9.5 = 5.7 , \text{km/s}$$
$$v_e = I_{sp} \cdot g_0 = 300 \cdot 9.81 = 2943 , \text{m/s}$$
Solve for $$\frac{m_0}{m_f}$$:
$$\frac{m_0}{m_f} = e^{\Delta v / v_e} = e^{5700 / 2943} \approx 5.5$$
Second Stage:
$$\Delta v_2 = 0.4 \cdot 9.5 = 3.8 , \text{km/s}$$
$$v_e = 450 \cdot 9.81 = 4414.5 , \text{m/s}$$
Solve for $$\frac{m_0}{m_f}$$:
$$\frac{m_0}{m_f} = e^{\Delta v / v_e} = e^{3800 / 4414.5} \approx 2.5$$
Total Mass:
Combine the stages iteratively to find the total initial mass $$m_0$$ that supports a 63,000 kg payload.
This is a simplified overview. Real-world calculations involve more detailed modeling of drag, gravity losses, and staging efficiencies. Tools like payload calculators or simulation software can refine these estimates.

SpaceNut · 2025-08-18 17:11:52

The standard landing legs for the first stage are roughly 500 kg each and are about 12 meters in length.

Other building blocks are the cygnus cargo ship.

AI Overview
Northrop Grumman's Cygnus Spacecraft: A Decade of Delivering ...
Cygnus spacecrafts have evolved from a standard configuration with ~2,750 kg capacity and 18 m³ volume to an enhanced version with ~3,750 kg and 27 m³, and now an upgraded "Mission B" with a capacity of 5,000 kg and a 36 m³ pressurized volume. These figures vary depending on the specific mission and version, with the enhanced and Mission B versions featuring longer pressurized modules to accommodate the increased cargo.
Evolution of Cygnus Cargo Specifications
Standard Version:
Payload Capacity: 2,750 kg
Pressurized Volume: 18 m³
Enhanced Configuration:
Payload Capacity: 3,750 kg
Pressurized Volume: 27 m³
Mission B (Upgraded Version):
Payload Capacity: 5,000 kg
Pressurized Volume: 36 m³
Key Details
What they carry:
Cygnus spacecrafts deliver food, water, spare parts, repairs, and scientific investigations to the International Space Station (ISS).
Design Evolution:
The spacecraft's design has been updated by Northrop Grumman to meet evolving customer needs, including increased cargo capacity.
Late Load Capability:
Enhanced and Mission B Cygnus versions allow for "late load" capability, meaning special equipment and science needs can be added shortly before launch.
Return Cargo:
Cygnus spacecrafts dispose of ISS waste by burning up upon atmospheric reentry.

he Northrop Grumman Cygnus cargo spacecraft typically carries around 800 kg of hypergolic propellant.
The fuel is stored in the Service Module and powers the main engine and smaller thrusters for navigation and orbit adjustments. Cygnus propellant specifications Propellant type:
The Cygnus uses hypergolic propellant, a mixture of hydrazine ($N_{2}H_{4}$) fuel and nitrogen tetroxide ($N_{2}O_{4}$) or MON-3 oxidizer. Hypergolic propellants ignite spontaneously upon contact, making them easy to use in space without a separate ignition system.
Propulsion system: The propulsion system features a main engine for major orbital adjustments and 32 smaller thrusters for attitude control.
Mission tasks: The fuel is used to perform various maneuvers throughout the mission, including:Phasing and rendezvous with the International Space Station (ISS).Occasional reboosts of the ISS to counteract atmospheric drag.A final burn to de-orbit the spacecraft for a destructive reentry into Earth's atmosphere.
Note on variants While the Standard Cygnus spacecraft had a similar fuel mass, information specifically points to the Enhanced version carrying roughly 800 kg of propellant. Later iterations, like those launched by the Falcon 9, have a larger overall launch mass and carry more propellant to support a heavier payload

The difference between the crew and cargo dragon is around 1,000kg with the cargo coming in dry at 6,600 kg mass.

Of course putting the pieces ordered in the correct order is the key to making the mission even possible.

SpaceNut · 2025-08-20 14:41:04

Falcon 9 has a diameter of 3.6 m and 70 m tall so its volume is V=πr2h=π·1.82·70≈712.51321 m³

So if we go for a habitat that is 10 m in diameter then we have 10 m tall to get V=πr2h=π·52·10≈785.39816 m³

So the question is the tank mass and engines to land something with the dry mass of a falcon.

post 51 has the numbers for the habitat with basically a second stage makes for more homework.

LEO to LLO rocket equation for the earth departure staging.

To calculate the delta-v required for a transfer from Low Earth Orbit (LEO) to Low Lunar Orbit (LLO), the Tsiolkovsky rocket equation is used in conjunction with orbital mechanics principles. Here's a concise breakdown:
1. Tsiolkovsky Rocket Equation
The equation is:
$$\Delta v = I_{sp} \cdot g_0 \cdot \ln\left(\frac{m_0}{m_f}\right)$$
Where:
Δv: Change in velocity (m/s)
Iₛₚ: Specific impulse of the rocket engine (s)
g₀: Standard gravity (9.81 m/s²)
m₀: Initial (fueled) mass of the rocket
mₓ: Final (dry) mass of the rocket

2. Delta-v Budget for LEO to LLO
The transfer typically involves:
LEO to Trans-Lunar Injection (TLI): ~3.2 km/s
Lunar Orbit Insertion (LOI): ~0.8 km/s
Thus, the total Δv is approximately 4.0 km/s.
3. Application of the Rocket Equation
Using the total Δv (4.0 km/s), you can calculate the required fuel mass ratio:
$$\ln\left(\frac{m_0}{m_f}\right) = \frac{\Delta v}{I_{sp} \cdot g_0}$$
Rearranging:
$$\frac{m_0}{m_f} = e^{\frac{\Delta v}{I_{sp} \cdot g_0}}$$
Plug in the values for Δv, Iₛₚ, and g₀ to determine the mass ratio.
Example
For a rocket with Iₛₚ = 450 s:
$$\ln\left(\frac{m_0}{m_f}\right) = \frac{4000}{450 \cdot 9.81} \approx 0.91$$
$$\frac{m_0}{m_f} = e^{0.91} \approx 2.48$$
This means the rocket's initial mass must be 2.48 times its dry mass to achieve the transfer.
This calculation assumes ideal conditions and does not account for inefficiencies, gravity losses, or trajectory corrections.

So the landing from orbit propulsive is

The propulsive landing of a rocket involves using its engines to decelerate and control its descent, ensuring a safe landing. This process is governed by physics principles, including the Tsiolkovsky rocket equation and Newton's laws of motion. Here's a simplified explanation of the key equation used in such scenarios:
Key Equation for Propulsive Landing
The velocity change (Δv) required for a rocket to decelerate and land safely is derived from the Tsiolkovsky rocket equation:
$$ \Delta v = v_e \ln\left(\frac{m_0}{m_f}\right) $$
Where:
Δv: Change in velocity needed for landing.
v_e: Effective exhaust velocity (specific impulse × gravitational acceleration).
m_0: Initial mass of the rocket (including remaining fuel).
m_f: Final mass of the rocket (after burning fuel).
Additional Considerations for Landing
Thrust-to-Weight Ratio: The rocket must generate enough thrust to counteract gravity and decelerate. The thrust-to-weight ratio is given by: $$ T/W = \frac{T}{m \cdot g} $$ Where:
T: Thrust produced by the engines.
m: Current mass of the rocket.
g: Gravitational acceleration (≈9.81 m/s² on Earth).
Burn Time: The duration of the engine burn to achieve the required Δv: $$ t = \frac{m_0 - m_f}{\dot{m}} $$ Where:
t: Burn time.
\dot{m}: Mass flow rate of the propellant.
Controlled Descent: The rocket must balance thrust and gravity to achieve a soft landing. This involves precise control of the throttle and orientation.
Practical Application
For a successful landing, engineers simulate and calculate:
The required Δv based on the rocket's altitude, velocity, and mass.
The optimal burn time and thrust profile to ensure a smooth deceleration.
Real-time adjustments using onboard sensors and guidance systems.
This combination of physics and engineering ensures rockets like SpaceX's Falcon 9 can land safely and be reused!

SpaceNut · 2025-08-22 14:50:52

Development of a Lunar Lander Modeling and Simulation Capability

The propulsive landing of a rocket involves using its engines to decelerate and control its descent, ensuring a safe landing. This process is governed by physics principles, including the Tsiolkovsky rocket equation and Newton's laws of motion. Here's a simplified explanation of the key equation used in such scenarios:
Key Equation for Propulsive Landing
The velocity change (Δv) required for a rocket to decelerate and land safely is derived from the Tsiolkovsky rocket equation:
$$ \Delta v = v_e \ln\left(\frac{m_0}{m_f}\right) $$
Where:
Δv: Change in velocity needed for landing.
v_e: Effective exhaust velocity (specific impulse × gravitational acceleration).
m_0: Initial mass of the rocket (including remaining fuel).
m_f: Final mass of the rocket (after burning fuel).

Additional Considerations for Landing

Thrust-to-Weight Ratio: The rocket must generate enough thrust to counteract gravity and decelerate. The thrust-to-weight ratio is given by:
$$ T/W = \frac{T}{m \cdot g} $$
Where:
T: Thrust produced by the engines.
m: Current mass of the rocket.
g: Gravitational acceleration (≈9.81 m/s² on Earth).

Burn Time: The duration of the engine burn to achieve the required Δv:
$$ t = \frac{m_0 - m_f}{\dot{m}} $$
Where:
t: Burn time.
\dot{m}: Mass flow rate of the propellant.

Controlled Descent: The rocket must balance thrust and gravity to achieve a soft landing. This involves precise control of the throttle and orientation.

Practical Application
For a successful landing, engineers simulate and calculate:
The required Δv based on the rocket's altitude, velocity, and mass.
The optimal burn time and thrust profile to ensure a smooth deceleration.
Real-time adjustments using onboard sensors and guidance systems.
This combination of physics and engineering ensures rockets like SpaceX's Falcon 9 can land safely and be reused!
There is no single "LLO 20,000 kg moon landing equation"; instead, a complex set of equations and principles govern a moon landing, including the Tsiolkovsky rocket equation to determine propellant requirements and orbital mechanics to calculate trajectories for rendezvous and descent. The Apollo Lunar Module, weighing approximately 15,200 kg, used these principles for its lunar orbit and descent, which would not be a simple formula.
Key Concepts and Equations
Tsiolkovsky Rocket Equation:
This fundamental equation relates the change in velocity (delta-v) of a rocket to its exhaust velocity and the ratio of its initial to final mass.
Δv = Ve * ln(m0 / mf)
Δv is the change in velocity.
Ve is the exhaust velocity.
ln is the natural logarithm.
m0 is the initial mass of the spacecraft.
mf is the final mass (after fuel is expended).
Orbital Mechanics:
Complex equations are used to predict and control the spacecraft's path in orbit around the Earth and Moon, ensuring a successful rendezvous and landing.
How it Applies to a Moon Landing
1. Mission Planning:
Engineers use these equations to determine how much fuel is needed for a lunar module with a specific payload, such as a 20,000 kg lander, to reach the Moon, enter lunar orbit, and then land safely.
2. Descent & Landing:
The equations calculate the necessary adjustments in velocity and thrust to slow the lunar module from orbital speed to a soft landing on the Moon's surface.
3. Ascent & Return:
The same principles are applied for the ascent to rejoin the command module and the return journey to Earth.
Example of Mass
The Apollo Lunar Module (LEM) had a standard launch mass of approximately 15,200 kg (33,500 lb).

AI Overview To determine the engine thrust required for a 20,000 kg lunar landing, you would use Newton's Second Law of Motion. The required thrust must be greater than the gravitational force of the Moon acting on the lander to allow for a controlled descent. A simplified calculation is provided below, followed by more complex considerations. Equation for required thrust The most basic equation for the thrust ($F_{T}$) required for a controlled lunar descent is:$F_{T}>m\times g_{L}$ Where: $F_{T}$ is the engine thrust (in Newtons)$m$ is the mass of the lander (in kg)$g_{L}$ is the acceleration due to gravity on the Moon, which is approximately 1.62 m/s². For a 20,000 kg lander, the minimum thrust would be:$F_{T}>20,000\text{\ kg}\times 1.62\text{\ m/s}^{2}$$F_{T}>32,400\text{\ N}$ or 32.4 kN Key considerations and complexities Controlled descent: This equation calculates the minimum thrust needed to hover. For a controlled landing, the engine must be "throttleable"—meaning its thrust can be adjusted. To slow the descent and land softly, the engine's thrust must be greater than gravity. The Apollo Lunar Module, for example, had a descent engine with a thrust that could be varied, or "deep throttled," over a wide range.Mass is not constant: A rocket's total mass decreases as it burns propellant. This means less thrust is needed to counteract gravity as the spacecraft empties its fuel. Modern landers must be designed to accommodate this change in mass and the resulting change in thrust requirements.Total "delta-v" (ΔV): The total change in velocity required for the mission is a more comprehensive metric than just the instantaneous thrust. The Tsiolkovsky Rocket Equation calculates the total ΔV capability of a spacecraft based on its propellant and engine efficiency. A lunar landing requires a specific ΔV budget, typically around 2,050 m/s for the descent phase.Thrust-to-weight ratio (TWR): Engineers use the TWR to describe engine performance relative to the spacecraft's mass. The TWR is typically calculated relative to Earth's gravity but can also be expressed for a lunar landing. For a lunar descent, an optimal TWR at the start of the burn might be in the range of 0.35 to 0.5 (using Earth weight)

SpaceNut · 2025-08-22 14:56:18

AI Overview
Landing a 12,000 kg payload on the Moon would require an estimated 22,000 kg of fuel, assuming a modern cargo lander design starting from a lunar orbit. This is based on calculations for a theoretical lander designed to maximize payload. The total mass of the lander before descent would be approximately 42,300 kg, which includes fuel, payload, and the dry mass of the lander itself.
Factors influencing the fuel estimate
The exact amount of fuel depends on several variables, including the lander's design, engine performance, and mission profile.
Delta-v: This measure of the change in velocity is the most important factor in calculating propellant needs.
Low Lunar Orbit (LLO) to surface: Landing from LLO requires a delta-v of about 1,720 m/s. For context, the Apollo Lunar Module descent stage was budgeted for 2,125 m/s, demonstrating that extra fuel is needed for contingency and precise maneuvers.
Specific Impulse (Isp): This measures how efficiently a rocket engine uses propellant. A higher Isp means more thrust is generated per unit of fuel, reducing the amount of propellant needed.
LOX/LH2 engines: Using liquid hydrogen and liquid oxygen offers a higher Isp than older, storable propellants. Some modern lander concepts plan to use these, as demonstrated by the Starship Human Landing System (HLS) for the Artemis program.
Mass ratio: This is the ratio of the fully fueled vehicle's mass to its dry mass (structure and payload).
A study examining a potential cargo lunar lander with a 12,000 kg payload projected a wet-to-dry mass ratio of over 4:1 for the descent stage.
Propellant type:
LOX/Methane: SpaceX's Starship HLS, for example, is designed to be refueled in Earth orbit and uses liquid oxygen and liquid methane. This allows it to deliver a large payload to the lunar surface.
In-Situ Resource Utilization (ISRU): Future missions plan to harvest lunar-derived propellant (LDP) from the Moon's surface. Since oxygen makes up 75% of the propellant mass for LOX/methane engines, this approach would significantly reduce the fuel that needs to be carried from Earth.
Comparison with past missions
Apollo Lunar Module (LM): The Apollo 11 LM's descent stage had a wet mass of 15,200 kg, carrying over 8,000 kg of hypergolic propellant to land a combined payload and dry mass of about 7,000 kg. This was a less efficient system with a smaller payload than the one you specified.
Summary of an example mission
A 2015 study for a modern cargo lander provides the best available estimate for a 12,000 kg payload:
Payload mass: 12,000 kg
Lander dry mass: 8,300 kg
Propellant mass: 22,000 kg
Total mass before descent: 42,300 kg

To transport a 42,300 kg vehicle from Low Earth Orbit (LEO) to Low Lunar Orbit (LLO), a significant amount of fuel is required, but the precise amount is not a single number. The required fuel mass depends on the engine's efficiency (specific impulse) and the mission profile chosen.
Assumptions for this estimate
To calculate the approximate fuel needed for a 42,300 kg departure from LEO, we must rely on a set of standard assumptions:
Total Delta-V (ΔV): The total change in velocity required to go from LEO to LLO is approximately 4.04 km/s (4,040 m/s). This is an aggregate of several maneuvers, including Trans-Lunar Injection (TLI), mid-course corrections, and Lunar Orbit Insertion (LOI).
Engine Type: The calculation uses performance figures for a high-performance chemical rocket, such as a liquid hydrogen/liquid oxygen (LH2/LOX) engine. This type of engine was used for the upper stages of the Saturn V and other missions.
Specific Impulse (Isp): A typical vacuum Isp for an LH2/LOX engine is around 450 seconds.
Fuel calculation
The calculation for fuel mass uses the Tsiolkovsky Rocket Equation:
ΔV = g₀ * Isp * ln(M₀ / Mf)
Where:
ΔV = Change in velocity (4,040 m/s)
g₀ = Standard gravity (9.807 m/s²)
Isp = Specific impulse (450 s)
M₀ = Initial mass (42,300 kg + fuel)
Mf = Final mass (42,300 kg)
Rearranging the equation to solve for the initial mass (M₀):
M₀ = Mf * e^(ΔV / (g₀ * Isp))
Step 1: Calculate the mass ratio
Mass Ratio = e^(4040 / (9.807 * 450))
Mass Ratio = e^(0.916) ≈ 2.50
Step 2: Calculate the initial mass (M₀) and fuel mass
M₀ = 42,300 kg * 2.50 ≈ 105,750 kg
Fuel Mass = M₀ - Mf
Fuel Mass = 105,750 kg - 42,300 kg ≈ 63,450 kg
Contextual considerations for the result
The calculation for 63,450 kg of fuel provides a simplified estimate. In practice, the actual fuel required can be higher due to several factors:
Propellant Reserve: Propellant is always needed for attitude control, mid-course corrections, and landing attempts (in the case of a lunar lander). This was not factored into the basic calculation.
Mission Profile: Missions can be optimized for fuel efficiency at the cost of longer travel times. For example, a low-energy transfer uses less fuel but can take months instead of days.
Spacecraft Architecture: A multi-stage architecture is much more efficient than a single-stage design. The calculation assumes a single, high-performance propulsion stage.
Alternative Propulsion: Advanced systems like electric propulsion offer much higher specific impulses, but with very low thrust. This would be unsuitable for the rapid burn-time maneuvers necessary for a quick Hohmann transfer

SpaceNut · 2025-08-22 15:00:50

Now to work number launching from the lunar surface to LLO.

Lunar launch mass is highly variable, depending on the payload and mission, but a typical ratio can be seen with the Apollo Lunar Module, where its total mass of 15,103 kg included 2,376 kg of propellant in the ascent stage and more in the descent stage. Modern Starship designs aim to bring significantly larger payloads, with a lunar variant requiring approximately 1,500 tons of propellant.
Apollo Lunar Module (Example)
Total Mass: 15,103 kg (including fuel and oxidizer)
Ascent Stage Propellant: 2,376 kg
Descent Stage Propellant: 8,200 kg (DPS propellants)
Lunar Lander Engines: These are specialized engines designed for the precise maneuvers required for landing on and taking off from the Moon's surface. Examples include:
Apollo Lunar Module Descent Engine: Developed by Thompson Ramo Wooldridge Inc (TRW), this engine could be throttled between 1,000 and 10,000 pounds of thrust, making it the most powerful LM engine. It was crucial for controlled landings and could be throttled for hovering and precise site selection.
Apollo Lunar Module Ascent Engine (LMAE): Built by Bell Aerosystems and later redesigned by Rocketdyne, this engine lifted the ascent stage off the Moon's surface and rendezvoused with the orbiting Command Module. It produced 3,500 pounds of thrust.
Reaction Control System (RCS) Thrusters: These small thrusters are used for attitude control and minor orbital adjustments. The Apollo LM had sixteen RCS thrusters, each with 100 pounds-force (440 N).

Here's a breakdown of how the figure relates to the Apollo program:
LLO Departure Vehicle: In the context of the Apollo missions, the LM was the vehicle that departed from Low Lunar Orbit (LLO) to land on the moon. The term "LLO departure vehicle" refers to this stage of the mission.
Total propellant: The Apollo 11 mission's LM had a total propellant load of over 23,000 kg, or approximately 53,000 pounds.
Fuel vs. Propellant: The LM used hypergolic propellants, meaning the fuel (Aerozine 50) and oxidizer (nitrogen tetroxide) ignited on contact, without a separate ignition system.
Propellant breakdown for a typical Apollo LM
The actual mass of propellant was divided between the two stages of the LM. While the exact figures varied by mission, for Apollo 11 the propellant breakdown was:
Descent stage propellant: 18,184 lbs (approx. 8,250 kg). This was used to descend from LLO to the lunar surface.
Ascent stage propellant: 5,238 lbs (approx. 2,376 kg). This was used to launch off the lunar surface and rendezvous with the Command and Service Module (CSM) in lunar orbit

SpaceNut · 2025-08-22 15:28:04

Lunar Ascent and Rendezvous Trajectory Design

The Apollo Lunar Orbit Rendezvous Architecture Decision Revisited

Analysis of Alternative Architectures for a 2024 Lunar Sortie

Just trying to get real numbers for each piece.

The Apollo Lunar Module's descent stage specifications varied by mission series, but generally, the H-series descent stage had a dry mass of around 2,034 kg (4,485 lbs) and a propellant mass of about 8,248 kg (18,184 lbs), with a total mass (including propellant) of approximately 10,282 kg (22,669 lbs).
The J-series lunar modules had a total descent stage mass of about 11,665 kg (25,718 lbs). These figures represent the weight of the lower, octagonal-shaped section of the LM that served as its platform for landing on the Moon, from which the ascent stage later launched.
Here are some key specifications:
Descent Stage Dry Mass: Approximately 2,034 kg (4,485 lbs).
Propellant Mass: Around 8,248 kg (18,184 lbs).
Total Mass (with propellant): Approximately 10,282 kg (22,669 lbs) for the H-series.
Total Mass (with propellant) for J-series: Around 11,665 kg (25,718 lbs).
Propellant Type: Hypergolic bipropellants, specifically nitrogen tetroxide (NTO) and Aerozine-50 (a mixture of hydrazine and UDMH).
Dimensions: The descent stage was roughly an octagonal prism, about 4.2 meters across and 1.7 meters thick, with a total landing gear span of 9.4 meters.
Descent Engine: A throttled engine (TR-201) designed by TRW, capable of thrust from roughly 4.67 kN to 45.04 kN (1,050 to 10,125 lbf).
The descent stage was the foundation for the lunar landing and housed the landing gear, the descent engine, and other critical systems, while the ascent stage was the crew cabin that would later return them to the Command Module

The Apollo Lunar Module (LM) ascent stage had a total mass of approximately 4,547 kg (9,987 lb) to 4,780 kg (10,538 lb), with a dry mass of around 2,383 kg (5,254 lb) and 2,358 kg (5,199 lb) of propellant, which was enough to launch the crew from the Moon and rendezvous with the command module.
Mass Breakdown
Total Mass: Approximately 4,547 kg (9,987 lb) to 4,780 kg (10,538 lb)
Dry Mass (without propellant): Around 2,383 kg (5,254 lb)
Propellant Mass: About 2,358 kg (5,199 lb) to 2,375 kg (5,236 lb)
Key Features
Purpose: The ascent stage was the upper part of the LM, containing the cabin, controls, and the engine for returning the astronauts to orbit.
Propellant: The ascent engine used NTO/Aerozine-50 propellant.
Crew: The stage was designed for a crew of two astronauts.
Weight Savings: To save weight, the LM had a minimal structure, with large structural members machined from single pieces to reduce the number of joints and fasteners

SpaceNut · 2025-08-23 19:17:30

Lots of number to tally for a mission

The Lunar Module's crew mass is approximately 144 kg (317 lb), representing the combined weight of two astronauts in their suits and equipment. This crew mass is carried by the Lunar Module's two-person ascent stage, which was designed for space operations on and around the Moon.
Key Specifications Related to Crew Mass
Crew Capacity: 2
Total Mass (Ascent Stage, Loaded): Approximately 4,780 kg (10,538 lb)
Crew Mass (Approximate): 144 kg (317 lb)
Pressurized Volume: 6.7 m³ (235 cu ft)
Habitable Volume: 4.5 m³ (160 cu ft)
The crew mass is an important component that influences the overall mass of the spacecraft and is factored into its performance and mission design.

During the Apollo missions, the two-person lunar module (LM) crew was provisioned with a food supply that weighed approximately 0.77 kg (1.7 lbs) per man per day, including packaging. The extremely small weight and volume envelope of the spacecraft required that most foods be dehydrated.
Food system specifications
Daily caloric intake: Each crew member received approximately 2,800 calories per day, divided into three meals.
Dehydrated food: About 80% of the food weight was removed by dehydrating it, with water from the LM's fuel cells used to rehydrate the meals during the mission. The freeze-drying process made the food extremely light and compact, with an almost indefinite shelf life.
Packaging: Meals were packed in four-ply, laminated film pouches to protect against moisture and oxygen invasion. Color-coded Velcro patches were attached to meal packages for easy identification.
Food preparation: Unlike the Command Module, the Lunar Module had no hot water dispenser, so food consumed on the lunar surface was rehydrated with cold water.
Bite-sized cubes: Astronauts also ate ready-to-eat, bite-sized food cubes coated with gelatin to prevent crumbs, which could interfere with sensitive equipment.
Snacks: Some missions included high-nutrient food bars, which astronauts could nibble on while wearing their helmets during moonwalks

The Apollo Lunar Module's crew oxygen supply varied by mission, with early missions using less than 100 pounds of gaseous oxygen for lunar descent and surface operations. The supply was separated into tanks in the descent stage for surface activities and tanks in the ascent stage for the return to the command module. Later, "J-missions" were equipped with additional capacity for longer stays.
Oxygen system mass specifications
Apollo 11 (Early mission)
Descent stage: The LM descent stage housed two oxygen tanks, each containing 48 pounds of gaseous oxygen at 2,690 psia. This supply covered descent, the surface stay, EVA backpack refills, and cabin repressurizations.
Ascent stage: For the rendezvous and docking after lunar liftoff, the ascent stage had two tanks, each containing 2.43 pounds of oxygen at 840 psia.
Extended "J-missions" (e.g., Apollo 15, 16, 17)
The final three Apollo missions carried greater oxygen capacity to support longer stays on the lunar surface. While the specific additional mass is not detailed in sources, the increased duration from a few hours (Apollo 11) to nearly three days (Apollo 17) necessitated a larger oxygen reserve.
Oxygen system context
The LM oxygen system was designed to support the astronauts for a limited time away from the Command and Service Module (CSM), which housed the primary liquid oxygen supply for the journey to and from the Moon.
Oxygen use
Cabin pressurization: The LM's atmosphere was pure oxygen at a pressure of 4.8 pounds per square inch.
Suit repressurization: Oxygen was used to repressurize the cabin and refill the Portable Life Support Systems (PLSS) worn by the astronauts during their moonwalks.
Metabolic consumption: The crew consumed oxygen for breathing. For instance, the descent tank on early missions was sized to allow for six PLSS refills and four cabin repressurizations.
Mass summary
Component Mass (pounds) Details Reference
Descent Stage Tank (x2) 96 Each tank held 48 lbs of gaseous oxygen.
Ascent Stage Tank (x2) 4.86 Each tank held 2.43 lbs of oxygen for the trip back to the CSM.

SpaceNut · 2025-08-24 13:15:03

Here is what Bezo is still working on for a lunar lander

Blue Origin's Blue Moon landers have two main configurations: the smaller, uncrewed Mark 1 and the larger, crewed Mark II for Artemis missions.
The Mark 1, a cylindrical uncrewed lander, is 8.05 meters tall, 3.08 meters in diameter, and can carry about 3 metric tons to the lunar surface using its BE-7 engine.
The Mark II is a larger, reusable human landing system capable of supporting four astronauts and delivering significant payloads to the lunar surface.
Mark 1 (Uncrewed Cargo Lander)
Purpose: Deliver scientific payloads and cargo to the lunar surface.
Dimensions: 8.05 meters tall, 3.08 meters in diameter.
Payload Capacity: ~3 metric tons (6,600 lb).
Propulsion: Powered by a single Blue Origin BE-7 engine using liquid hydrogen and liquid oxygen.
Power: Solar panels and fuel cells.
Navigation: Features an inertial measuring unit (IMU), a star tracker, and hot and cold gas thrusters.
Mark II (Human Landing System)
Purpose: Transport astronauts to the lunar surface as part of the Artemis program.
Dimensions: Around 16 meters tall.
Crew Capacity: Can accommodate four astronauts.
Payload Capacity: Up to 20 metric tons (reusable) or 30 metric tons (one-way).
Propulsion: Uses multiple BE-7 engines for deep throttling thrust.
Features: Includes a habitation module for astronauts, an airlock, and radiators for heat management.
Reusability: Designed to be reusable, allowing for return flights to Earth or refueling in lunar orbit.
Fuel: Powered by liquid hydrogen and liquid oxygen, with the capability to be refueled on the Moon.
Common Components & Technologies
BE-7 Engine:
.
A key component for both landers, designed for lunar applications with a 40 kN thrust.
New Glenn Rocket:
.
The landers are designed to be launched on Blue Origin's New Glenn rocket, fitting within its payload fairing.
Lunar Refueling:
.
Blue Origin is developing technology to store cryogenic propellants, like liquid hydrogen, on the lunar surface and in orbit for refueling.

https://www.blueorigin.com/blue-moon/mark-1

https://spaceinsider.tech/2023/09/28/blue-origins-hls/

SpaceNut · 2025-08-24 15:48:26

https://www.thenewatlantis.com/publications/moon-direct

https://www.marssociety.de/files/Medien … DC2018.pdf

SpaceNut · 2025-08-27 15:33:06

The main issue for rocket design is the combination for an engine ISP that we can purchase to build with.

Common high-performance rocket fuel and oxidizer combinations include Liquid Hydrogen (LH2) / Liquid Oxygen (LOX) for maximum specific impulse (Isp), and Liquid Methane (CH4) (Methalox) / LOX, which offers a good balance of Isp and storability. For simplicity and ignition, hypergolic propellants like Aerozine 50/NTO ignite on contact but are toxic. Less common choices include RP-1 (kerosene)/LOX for high density and HTPB/N2O for solid rocket motors.
This video explains the basics of rocket fuels and specific impulse:
56s
https://www.youtube.com/watch?v=95ROPl-ZObk&t=544
purdueMET
YouTube · Mar 26, 2020
High-Specific Impulse Combinations
Liquid Hydrogen (LH2) / Liquid Oxygen (LOX): Offers the highest specific impulse due to the low molecular mass of its exhaust gases. This is the most efficient chemical combination, making it ideal for high-performance upper stages, but LH2 requires massive tanks due to its low density.
This video discusses different types of rocket fuels and their properties:
56s
https://www.youtube.com/watch?v=jI8TuufCp0M&t=711
Scott Manley
YouTube · Apr 9, 2013
Balanced Combinations
Liquid Methane (CH4) / Liquid Oxygen (LOX) (Methalox):
.
Offers a good compromise between specific impulse and density. Methane burns cleaner than kerosene, is more efficient, and is easier to produce on other celestial bodies, making it a popular choice for reusable rockets like SpaceX's Starship.
RP-1 (Kerosene) / Liquid Oxygen (LOX):
.
A traditional combination that provides high thrust and density, leading to smaller propellant tanks compared to LH2. However, it has a lower specific impulse than cryogenics.
Hypergolic Combinations
Aerozine 50/Nitrogen Tetroxide (NTO): and Monomethylhydrazine (MMH)/NTO: These fuels ignite upon contact, eliminating the need for an igniter and providing reliable ignition, which is valuable for upper stages and reaction control systems. They are highly toxic and less efficient than other options, according to Wikipedia.com.
Solid Rocket Motor Propellants
HTPB/Nitrous Oxide (N2O): A combination for solid rocket motors, where the fuel (HTPB) and oxidizer (N2O) are combined in a solid matrix.
This video demonstrates how to make rocket fuel from household items:
55s
https://www.youtube.com/watch?v=TSzgvP0 … N5tD&t=170
Thoisoi2 - Chemical Experiments!
YouTube · Jan 22, 2022
Key Considerations for Propellant Choice
Specific Impulse (Isp): A measure of engine efficiency, with higher Isp indicating more thrust for the same amount of propellant over time.
Density: Determines the volume of tanks needed.
Storability: How easy a propellant is to store, with cryogens like LH2 requiring complex insulation and management.
Application: The overall mission requirements dictate the best combination

SpaceNut · 2025-08-27 15:38:14

first video Rocket engine performance (specific impulse) table

https://www.funkystuff.org/wp-content/u … mpulse.png

https://eaglepubs.erau.edu/app/uploads/ … 68x455.png

https://blogger.googleusercontent.com/i … able_1.png

SpaceNut · 2025-08-27 16:44:01

A few posts that might make a library of rocket engines to fuel information follows.

The SpaceX Falcon 9 rocket is fueled by a rocket-grade kerosene (RP-1) and liquid oxygen (LOX) propellant, which is mixed and ignited in its Merlin engines. Nine Merlin engines power the first stage for launch and stage separation, while a single, larger Merlin Vacuum engine powers the second stage. To achieve ignition, a pyrophoric mixture called triethylaluminum-triethylborane (TEA-TEB) is injected into the combustion chamber.

Propellants and Engine Types
RP-1: A highly refined kerosene used as the fuel.
Liquid Oxygen (LOX): The oxidizer, which is subcooled to enhance density and performance.
Merlin Engines: The family of engines developed by SpaceX.
First Stage: Powered by nine Merlin engines.
Second Stage: Powered by a single Merlin Vacuum engine, designed for the vacuum of space and capable of multiple restarts.
Engine Ignition
Propellant Injection: RP-1 and LOX are delivered to the combustion chamber.
Pyrophoric Ignition: A small amount of TEA-TEB is injected.
Spontaneous Ignition: This mixture ignites spontaneously upon contact with the oxygen.
Stable Burn: The initial combustion provides a reliable ignition source for the main RP-1/LOX mixture.
Why This System Works
High Thrust: The combination of RP-1 and LOX in the Merlin engines produces significant thrust for liftoff.
Reusability: The Merlin engines were designed for recovery and reuse.
Restart Capability: The TEA-TEB ignition system allows for multiple engine restarts, which is crucial for precise orbital maneuvers and stage separation.

The Falcon 9 rocket's second stage uses liquid oxygen (LOX) and rocket-grade kerosene (RP-1) as its propellants, which are stored in tanks made from an aluminum lithium alloy. A single Merlin Vacuum Engine on the second stage ignites after separation from the first stage and can be restarted multiple times to place payloads into different orbits.
Fuel and Oxidizer
Liquid Oxygen (LOX): This is the oxidizer, providing the oxygen needed for the engine to burn the fuel.
RP-1 (Kerosene): This is the fuel.
How it Works
1. Ignition:
After the first stage separates, a few seconds later, the second stage's single Merlin Vacuum Engine ignites.
2. Multi-Restart Capability:
The second stage can restart its engine multiple times, allowing it to deploy various payloads into different orbits during a single mission.
3. Propellant Delivery:
The propellants are fed from their respective tanks through plumbing into the engine's combustion chamber.
4. Combustion:
The liquid oxygen and RP-1 burn at high pressure and temperature, generating a massive amount of energy.
5. Payload Deployment:
After achieving the desired orbit, the engine cuts off, and the second stage deploys its payload.
Key Features
Propellant Mixture:
.
Like the first stage, the second stage uses liquid oxygen and RP-1 as propellants.
Vacuum Engine:
.
The second stage uses a specially designed Merlin Vacuum Engine, optimized for operation in the vacuum of space.
Cold Gas System:
.
A cold nitrogen gas attitude control system is used for precise control and orientation of the rocket in space, ensuring accurate payload placement.

SpaceNut · 2025-08-27 16:45:37

The first stage of the Vulcan rocket uses liquefied natural gas (methane) and liquid oxygen as fuel, powered by two Blue Origin BE-4 engines, while its Centaur V upper stage uses liquid hydrogen and liquid oxygen for its Aerojet Rocketdyne RL10 engines. Solid rocket boosters can also be attached to the first stage for additional thrust at liftoff.
The Vulcan rocket's second stage, known as the Centaur V, is powered by two Aerojet Rocketdyne RL10 engines that run on liquid hydrogen and liquid oxygen fuel. These engines are designed to precisely place payloads into orbit by providing the energy for the final stages of the mission.
First Stage
Engines: Two Blue Origin BE-4 engines.
Fuel: Liquefied natural gas (methane) and liquid oxygen.
Purpose: To lift the rocket off the launch pad and provide power for the initial phase of flight.
Second Stage (Centaur V)
Engines: Two Aerojet Rocketdyne RL10 engines.
Fuel: Liquid hydrogen and liquid oxygen.
Purpose: To provide precise, restartable thrust to accurately place payloads into their intended orbits.
Solid Rocket Boosters (Optional)
Fuel: Graphite-epoxy composite.
Purpose: To provide extra thrust during liftoff for heavier payloads.

Key details about the second stage:
Name: Centaur V
Engines: Two Aerojet Rocketdyne RL10C engines
Fuel: Liquid hydrogen (LH₂) and liquid oxygen (LO₂)
Purpose: To provide the precise final orbital insertion and high-energy orbits needed for complex payloads.
Manufacturer: Aerojet Rocketdyne is the manufacturer of the RL10 engine.

SpaceNut · 2025-08-27 16:47:00

New Glenn uses a dual-propellant strategy: its first stage uses seven BE-4 engines powered by liquid oxygen and liquid methane (liquid natural gas), while the second stage uses two BE-3U engines fueled by liquid oxygen and liquid hydrogen. This combination allows the rocket to maximize thrust with the first stage and achieve high efficiency with the second stage for orbital flight.

First Stage: Thrust and Methane Power
Engines: Seven BE-4 engines.
Propellants: Liquid oxygen (LOX) and liquid natural gas (LNG), which is also known as liquid methane.
Purpose: To provide the massive thrust needed to lift the rocket off the pad and carry it through Earth's atmosphere.
Second Stage: Efficiency and Hydrogen Power
Engines: Two BE-3U engines.
Propellants: Liquid oxygen (LOX) and liquid hydrogen (LH2).
Purpose: The use of highly efficient liquid hydrogen allows the second stage to place payloads into precise orbits with greater efficiency in the vacuum of space.
Key Differences in Propellants
Methane vs. Hydrogen:
Methane is denser and provides higher thrust, which is ideal for the powerful first stage, while hydrogen is more efficient, making it a better choice for the upper stage where efficiency is critical for performance.
Oxygen-rich Staged Combustion:
The BE-4 engines use an oxygen-rich staged combustion cycle to improve performance.
Expander Cycle:
The BE-3U engines utilize an expander cycle, which is known for its efficiency.

The New Glenn rocket's second stage is powered by two BE-3U engines fueled with liquid hydrogen and liquid oxygen (LH2/LOX). These engines are optimized for vacuum operation, providing a large thrust and allowing for mission-specific performance, such as direct payload injection into high-energy orbits.
Engine Details
Name: BE-3U (Vacuum-optimized)
Fuel: Liquid hydrogen (LH2) and liquid oxygen (LOX)
Configuration: Two BE-3U engines on the second stage
Thrust: Together, they produce over 320,000 pounds of vacuum thrust
Features:
Large nozzle extensions for efficient expansion in a vacuum.
In-space restart capabilities for increased mission versatility.
Derived from the BE-3 engine used on the New Shepard suborbital rocket

SpaceNut · 2025-08-27 16:47:58

Atlas rocket stages use different fuels depending on the component: the main first stage (or core) uses liquid oxygen and RP-1 (kerosene). The upper stage, the Centaur, uses liquid oxygen and liquid hydrogen, which is a highly efficient propellant. Strap-on boosters, when used, are solid rocket motors containing a form of solid fuel.
First Stage (Booster)
Propellant: Liquid oxygen (LOX) and RP-1 (a refined kerosene).
Engine(s): A single, Russian-made RD-180 engine powers the first stage of the Atlas V. Older Atlas versions used various Rocketdyne engines.
Second Stage (Upper Stage)
Propellant: Liquid oxygen (LOX) and liquid hydrogen (LH2).
Engine(s): The Centaur upper stage uses one or two Aerojet Rocketdyne RL10 engines. This liquid-hydrogen fueled stage is known for its high efficiency and ability to be restarted in space.
Strap-on Solid Rocket Boosters (SRBs)
Propellant:
A solid fuel containing powdered aluminum and an oxidizer called ammonium perchlorate.
Function:
These are added to the Atlas V to provide additional liftoff thrust for heavier payloads, and can be configured in different numbers depending on the mission

The second stage engine for the Atlas rocket (specifically the Atlas V and its predecessors, known as the Centaur upper stage) is powered by a combination of liquid hydrogen (fuel) and liquid oxygen (oxidizer). These cryogenic propellants are burned in an Aerojet Rocketdyne RL10 engine, which is known for its high energy and restart capability.
Here's a breakdown:
Propellant:
The RL10 engine uses liquid hydrogen as its fuel and liquid oxygen as its oxidizer.
Engine Type:
The engine is the RL10, a high-energy, American-made engine manufactured by Aerojet Rocketdyne (formerly Pratt & Whitney).
Function:
This high-energy propellant combination is ideal for upper stages like the Centaur, allowing it to efficiently place payloads into orbit with high precision.
Handling:
Liquid hydrogen is a very high-energy fuel but also a cryogenic fluid with significant handling challenges

SpaceNut · 2025-08-27 16:48:50

The Delta rocket's second stage is powered by a restartable Aerojet AJ10-118K engine, which burns hypergolic propellants: Aerozine-50 (a fuel) and nitrogen tetroxide (N₂O₄, an oxidizer). These propellants react on contact, eliminating the need for a complex ignition system and allowing for multiple restarts during the flight to achieve different orbits or adjust trajectory.
Key details about the Delta second stage:
Engine: Aerojet AJ10-118K.
Propellants: Hypergolic mixture of Aerozine-50 and N₂O₄.
Restart capability: The engine could restart up to six times, allowing for multiple burns and fine-tuning of the rocket's trajectory.
Control: The engine could gimbal for pitch and yaw control, while nitrogen thrusters provided roll control during coast phases and powered flight.
Tanks: The tanks were made of stainless steel and pressurized with helium gas

Delta rocket stage engines used a variety of fuels depending on the stage and model, with earlier versions using RP-1 (kerosene) and liquid oxygen (LOX), or hypergolic propellants like UDMH and IRFNA, while later models, particularly the Delta IV, used liquid hydrogen (LH2) and LOX for their high-efficiency first stages. Solid rocket boosters, a common feature on many Delta variants, also use a solid fuel mixture of ammonium perchlorate, aluminum powder, and a rubber-like fuel.
First Stage Fuels
Early Delta Models (Delta I/II): Utilized RP-1 (kerosene) and liquid oxygen (LOX).
Delta IV: Employed liquid hydrogen (LH2) and liquid oxygen (LOX) for the large, powerful main rocket engines.
Second Stage Fuels
Early Delta Models (Thor-Delta/Delta I):
.
Used hypergolic propellants, a mix of unsymmetrical dimethylhydrazine (UDMH) and inhibited red fuming nitric acid (IRFNA).
Later Delta Models (Delta III/IV):
.
transitioned to a liquid oxygen/liquid hydrogen combination for improved safety and efficiency.
Third Stage (Optional) Fuels
Solid Boosters:
.
Solid-fueled motors were often added to stages for extra thrust, using a solid propellant mix containing ammonium perchlorate (oxidizer), aluminum powder (fuel), and polybutadiene (fuel binder).
Third Stage Motors:
.
The solid-fueled Altair motors used for some Delta versions burned solid propellant, similar to other solid rocket booster

SpaceNut · 2025-08-27 16:50:12

The Artemis rocket's RS-25 core stage engines are fueled by a liquid hydrogen (LH2) and liquid oxygen (LOX) propellant combination, chosen for its superior specific impulse. Liquid hydrogen serves as the primary fuel, while liquid oxygen acts as the oxidizer, providing the necessary thrust to power the Space Launch System (SLS) rocket.
Fueling Process
Liquid Hydrogen:
.
A large volume of liquid hydrogen is stored in a cryogenic tank within the SLS core stage and is cooled to extremely low temperatures of -423 degrees Fahrenheit.
Liquid Oxygen:
.
Liquid oxygen is also stored in a cryogenic tank, with a separate tank and line running around the hydrogen tank.
Turbopumps:
.
Due to the vastly different densities and flow rates of liquid hydrogen and liquid oxygen, the RS-25 engines use two separate turbopumps to manage each propellant.
Combustion:
.
Both the fuel and oxidizer are fed into the RS-25 engines to produce thrust for the Artemis missions.
Key Considerations
Cryogenic Challenges:
.
Liquid hydrogen's extremely low temperature makes it difficult to work with, requiring well-insulated tanks and careful handling to prevent leaks and maintain the liquid state.
Efficiency:
.
The LOX/LH2 combination is used because it offers the maximum possible energy for the fuel's weight, providing a highly efficient engine.
New Engines for Future Missions:
.
While the first Artemis missions use modified space shuttle RS-25 engines, NASA is now producing new RS-25 engines for future missions, starting with Artemis V, to ensure a continued supply for the SLS program

The second stage of the Artemis rocket, known as the Interim Cryogenic Propulsion Stage (ICPS), is fueled with liquid hydrogen (LH2) and liquid oxygen (LOX) to power its single RL10 engine, providing the necessary boost to send the Orion spacecraft towards the Moon. It also carries hydrazine for its Reaction Control System (RCS), used for orientation and stabilization in space.
Details of the ICPS (Second Stage):
Propellants:
The main propellant for the ICPS engine is a combination of liquid hydrogen (LH2) and liquid oxygen (LOX).
Engine:
The RL10 engine is used in this stage, with the Artemis 2 mission debuting a newer RL-10C2 model.
Purpose:
After the SLS core stage and Solid Rocket Boosters (SRBs) complete their burn and separate, the ICPS uses its RL10 engine to provide the Trans-Lunar Injection (TLI) burn, sending the Orion spacecraft out of Earth's orbit and toward the Moon.
Hydrazine:
In addition to its main engine fuel, the ICPS carries hydrazine for its Reaction Control System, which is used for small attitude adjustments and maneuvering in space.
Fueling Process:
1. Preparation:
The ICPS is transported to NASA's Kennedy Space Center and then to the Vehicle Assembly Building (VAB).
2. Fueling:
Technicians then fuel the ICPS with its required propellants, including liquid hydrogen, liquid oxygen, and hydrazine.
3. Integration:
The ICPS is then mated with the rest of the SLS rocket elements for launc

SpaceNut · 2025-08-27 16:50:38

The Delta IV Heavy rocket uses liquid hydrogen (LH2) as fuel and liquid oxygen (LOx) as oxidizer for its three main RS-68A engines, located on the common booster cores (first stage), and for the RL10 engine in its second stage. At liftoff, the three engines ignite, burning the hydrogen and oxygen to generate immense thrust, with the two side boosters jettisoning after they run out of fuel and the central booster continuing to fire before separating.

First Stage (Common Booster Cores)
Engines:
Each of the three Common Booster Cores (CBCs) is powered by an Aerojet Rocketdyne RS-68A engine.
Propellants:
These engines burn liquid hydrogen (LH2) as the fuel and liquid oxygen (LOx) as the oxidizer.
Function:
At liftoff, all three engines ignite to provide maximum thrust. The center core then throttles down to conserve fuel while the side boosters continue at full power. Once the side boosters are depleted, they separate from the rocket, and the central booster ramps back up to full thrust to carry the rocket through the thickest parts of the atmosphere.
Second Stage
Engine: The Delta IV Heavy's upper stage uses a single RL10 engine.
Propellants: Like the first stage, this engine also uses liquid hydrogen (LH2) and liquid oxygen (LOx).
Function: After the first stage boosters are jettisoned, the RL10 engine ignites for one or more burns to propel the payload into its final orbit.

The Delta IV Heavy's second stage, the Delta Cryogenic Second Stage (DCSS), is powered by a single Aerojet Rocketdyne RL10 engine, which uses liquid hydrogen (LH2) and liquid oxygen (LOX) as propellants to fuel the engine and place the payload into its final orbit. The DCSS consists of a large, cylindrical LH2 tank with the LOX tank suspended below it, from which the RL10 engine, featuring an extendable carbon-carbon nozzle for improved performance, is attached.
Engine and Propellants
Engine: A single, vacuum-optimized Aerojet Rocketdyne RL10 engine is used.
Propellants: The engine burns liquid hydrogen (LH2) as the fuel and liquid oxygen (LOX) as the oxidizer, which are cryogenic propellants.
Stage Architecture
Unique Design:
The DCSS features a unique two-tank architecture, where the liquid hydrogen tank forms the upper section and carries the payload launch loads.
LOX Tank and Engine:
The LOX tank is suspended from the LH2 tank and provides structural support for the engine.
Function
Orbital Insertion:
After the first stage separates, the second stage ignites to perform the final engine burns.
Precision Control:
The RL10 engine's ability to restart and adjust its thrust is crucial for fine-tuning the rocket's trajectory to achieve the precise orbit required for the payload.

SpaceNut · 2025-08-28 15:59:15

seems this is a choice

Wikipedia
https://en.wikipedia.org › wiki › RL10
Thrust (altitude): 24,750 lbf (110.1 kN) · Specific impulse: 465.5 seconds (4.565 km/s) · Engine weight, dry: 664 lb (301.2 kg) · Height: 163.5 in (4.14 m) ...
AI Overview

+7
The RL10 is a high-performance, liquid-fueled upper-stage rocket engine from Aerojet Rocketdyne that uses liquid hydrogen and liquid oxygen (LH2/LOX) propellants. Key specifications include thrusts around 25,000 pounds (110 kN) for models like the RL10B-2, very high specific impulses of over 465 seconds, and an expander cycle. The engine's performance is highly reliable and it has been continuously upgraded for missions like NASA's Space Launch System (SLS) and commercial launchers such as Atlas and Delta.
Key Specifications (RL10B-2 Model):
Thrust: 24,750 lbf (110 kN)
Specific Impulse (Isp): 465.5 seconds (4.565 km/s)
Engine Weight (dry): 664 lb (301.2 kg)
Propellants: Liquid Hydrogen (LH2) and Liquid Oxygen (LOX)
Cycle Type: Expander Cycle
Propellant Mixture Ratio: ~5.88 to 1 (Oxygen:Hydrogen by mass)
Key Features and Development
High Performance:
The RL10 is the world's highest-performance upper-stage rocket engine, known for its high efficiency thanks to its expander cycle and expandable nozzle.
Reliability:
It has a long history of reliable operation, with millions of accumulated hot-fire seconds and a demonstrated reliability of 0.9998.
Evolution:
The engine family has been continuously upgraded since its original design for the Centaur upper stage. Modern versions incorporate features like 3D printing and extendable nozzles for improved performance.
Applications:
The RL10 powers upper stages for various launch vehicles, including the Space Launch System (SLS) Interim Cryogenic Propulsion Stage (ICPS) and Exploration Upper Stage (EUS), as well as ULA's Atlas and Delta rockets and the Vulcan Centaur.
Manufacturing:
Recent advancements involve using 3D printing to create components from materials like copper, which enables complex geometries not possible with traditional methods.

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#51 2025-08-17 14:26:59

Re: Dr. Robert Zubrin Moon Direct: How to build a moonbase in four years

#52 2025-08-18 15:27:40

Re: Dr. Robert Zubrin Moon Direct: How to build a moonbase in four years

#53 2025-08-18 16:19:44

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Re: Dr. Robert Zubrin Moon Direct: How to build a moonbase in four years

#55 2025-08-18 17:11:52

Re: Dr. Robert Zubrin Moon Direct: How to build a moonbase in four years

#56 2025-08-20 14:41:04

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#57 2025-08-22 14:50:52

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#59 2025-08-22 15:00:50

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#60 2025-08-22 15:28:04

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#61 2025-08-23 19:17:30

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#62 2025-08-24 13:15:03

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#63 2025-08-24 15:48:26

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#64 2025-08-27 15:33:06

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#65 2025-08-27 15:38:14

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#66 2025-08-27 16:44:01

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#67 2025-08-27 16:45:37

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#68 2025-08-27 16:47:00

Re: Dr. Robert Zubrin Moon Direct: How to build a moonbase in four years

#69 2025-08-27 16:47:58

Re: Dr. Robert Zubrin Moon Direct: How to build a moonbase in four years

#70 2025-08-27 16:48:50

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#71 2025-08-27 16:50:12

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#72 2025-08-27 16:50:38

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#73 2025-08-28 15:59:15

Re: Dr. Robert Zubrin Moon Direct: How to build a moonbase in four years

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