Hybrid Ultimate Rocket 2

I would like to further enhance my Hybrid Ultimate Rocket design. The original concept featured a rocket-inside-a-rocket layout, which required complex support structures between the outer and inner stages. This meant more dead weight and mechanical complexity. To optimize this, I kept the design of alternating High-Test Peroxide (HTP) and Liquid Petroleum Gas (LPG) fuel tanks distributed as a circular ring. The upper stages repeat this same architecture. This allows for significantly lower and simpler structural support, easier stage separation, and one more bonus that I will explain later.

HTP requires gentle handling and a minimal flow path to the engines to reduce mechanical risks. I opted for allocating a unified aerospike engine module under each structural joint. The engines are directly fed by the adjacent tanks and mounted below the studs that connect the tanks. These studs extend to the top of each stage, providing a direct, straight-up path for thrust transfer. This eliminates unnecessary structural framing, reducing dead weight and simplifying the design.

Simplifying the piping required eliminating all fluid connections between the individual tanks. Rockets require continuously lower thrust over time due to reduced propellant weight and the need to limit the g-force applied to the structure. Therefore, each engine pair burns for a different duration, and by the time an engine is shut down, its attached tanks are completely depleted. To maintain a uniform height for all tanks across the stage, the diameter of the tanks is adjusted for each individual column. This gradual tank diameter change is perfectly balanced and distributed around the ring to minimize the pressure and volume differences between adjacent tanks.

Both the HTP and LPG will be fed using flexible PFA (Perfluoroalkoxy alkane) tubing. These tubes will be externally squished by stepper motors to control their flow, which is the safest way to throttle the HTP without internal valve cavities. The PFA tubing joins to SiO₂ (fused silica) tubing before interfacing with the Silicon Carbide (SiC) injector piping, creating a transiently thermally insulating, hermetically sealed setup.

The unified aerospike engine modules are sintered from solid SiC. The engines will not feature regenerative cooling channels, allowing for a simpler casting and sintering production process. The lower exhaust temperature of the HTP+LPG propellant chemistry, coupled with the 2200°C thermal capability of sintered SiC, negates the need for protective regenerative cooling. As a result, less pressure is dropped across the fuel lines and the soot problem is entirely mitigated. The engines operate at a low injector pressure of 6 bar, which requires minimizing any pressure losses on the way to the combustion chamber.

The tanks themselves are made of seamless, extruded PFA inner liners, wrapped externally by thin-gauge stainless steel. The vertical structural studs of the rocket are also made of stainless steel. This reduces thermal coupling from the hot engines up into the airframe, increases the structural strength of each module, and allows for straightforward automated laser welding and construction.

The final stage features a carbon-fiber-rod-supported fabric fairing on its nose. This fabric is semi-permeable, bleeding a small percentage of air into the internal void of the rocket structure. This continuous micro-bleed fills the base area, reducing the vacuum effect seen at higher altitudes and neutralizing the base drag spike natively. Due to its lightweight and simple design, this fabric cone is easily discarded at the initial stage separation altitude of 100 km. The near-zero separation speed in vacuum will make the separation of the first and upper stages seamless. While Stage 2 and Stage 3 separations occur at much higher hypersonic speeds, the complete lack of air further simplifies the structural release.

The first stage also features a matching fabric fairing structure on its upper rim, though this one is kept closed to the air during ascent. This structure becomes highly advantageous when the first stage separates and returns to the launch site. During descent, this bleeding fabric functions as a high-drag aerodynamic parachute, limiting the stage's terminal velocity before touchdown. This drastically reduces the propellant reserve needed for the final landing burn. Controllable shutters on the fabric allow the flight computer to modulate the air bleed, providing aerodynamic maneuvering and steering during landing. Due to the higher aerodynamic stresses experienced during atmospheric reentry, the first-stage fairing is structurally reinforced compared to the upper-stage fairing.

Additionally, this architecture preserves a massive hollow void space inside the core of the second and third stages of the rocket, reserved entirely for the payload. This allows high-volume, low-density payloads to be launched to orbit, such as monolithic space station modules and large telescopes. This huge empty space allows satellites to be deployed in their final, fully extended shapes without requiring complex folding mechanisms. Launching payloads unfolded increases their structural rigidity, lowers their weight, and eliminates the common failure modes associated with orbital unfolding procedures.

The structural design of the rocket requires all tank modules to possess their own dedicated engines to maintain direct load paths up the studs. If the first stage uses 64 engines, the second and third stages must also feature 64 engines. This means the individual engine dimensions scale down significantly for the upper stages, resulting in minuscule modules. Having such a high number of small, independent engines allows for gimbal-less differential thrust maneuvering and highly precise throttling profiles. This approach enables the use of cheaper, lighter, mass-produced ceramic components, making upper-stage manufacturing and scaling highly cost-efficient and enabling rapid launch turnaround times.