3.1 Option A (Second-Stage Configuration, 2.25m Diameter)
1) The first stage is a 2.25m module for a new-generation launch vehicle. The engine is upgraded from a single axis TVCto a dual-axis configuration, with a propellant mass of 61 tons and a 4x1000N hydrogen peroxide/kerosene auxiliary power system for roll control.
2) The second stage uses a 15t-class liquid oxygen-kerosene engine, with a dual-pendulum, double-start configuration and a propellant mass of 13.15 tons. Roll control, attitude control during ballistic cruise, and propellant management are controlled by 10x25N, 4x100N, and 2x300N hydrogen peroxide/kerosene auxiliary power systems.
3) The fairing has a diameter of 2.25m or 2.6m and can be hoisted in one piece or in separate pieces.
4) The rocket has an overall length of 35.07m and a diameter of 2.25m.
5) The takeoff mass is 84.537t, the takeoff thrust-to-weight ratio is 1.467, and the maximum dynamic pressure is 33.8 kPa, maximum axial overload 5.8 g;
6) 700 km SSO payload capacity approximately 650 kgg
3.2 Option B (three-stage configuration, 2.25 m diameter)
1) First stage same as Option A;
2) Second stage essentially the same as Option A, with single engine start;
3) Add a third stage, using 4x1000N trajectory maneuvers, 4x25N and 8x100N for second-stage roll control, and third-stage attitude control. Compared to Option A, 4 x 1000N thrust chambers and 4 x WON thrust chambers are added, while 6 x 25N thrust chambers and 2 x 300N thrust chambers are reduced, resulting in an increased fuel volume.
4) The fairing diameter is 2.25m or 2.6m, and can be hoisted in one piece or in separate pieces.
5) The rocket is 35.87m long and 2.25m in diameter.
6) The takeoff mass is 85.027t, the takeoff thrust-to-weight ratio is 1.459, the maximum dynamic pressure is 36.9kPa, and the maximum axial overload is 5.7g.
7) The 700km SSO payload capacity is approximately 870kg.
Comparison between Option B and Option A:
1) Adaptability: The three-stage option (Option B) significantly increases payload capacity (from 650kg to 870kg for the 700km SSO, a 34% increase; from 200kg to 670kg for the 1200km SSO, a 235% increase). See Figure 2. It also has the ability to leave orbit, which greatly increases its adaptability to different satellites.*
2) Complexity: The three-stage design adds a third stage, slightly increasing the complexity of the auxiliary power system and requiring one more separation. However, the YF-115 engine is switched to a single-stage start, simplifying the pressurization system, thus maintaining the same complexity.
3) Economical: With increased payload capacity, the rocket cost per unit payload decreases significantly. Therefore, small launch vehicles should focus on the three-stage design.
However, Plan B has the following issues:
1) The fineness ratio reaches 15.9, limiting the vehicle height and making it impossible to launch dual or larger satellites.
2) The takeoff thrust-to-weight ratio reaches 1.459, and the maximum dynamic pressure reaches 36.9 kPa, resulting in a relatively heavy payload.
These issues can be addressed by the following approaches:
1) Reducing the YF-100 engine thrust to 110 tons would resolve the maximum dynamic pressure issue, but would reduce the payload capacity by approximately 130 kg (see Figure 3).
2) Other approaches: Optimizing the tank bulkhead shape, appropriately lowering the engine height, integrating the instrument bay with the payload support, and structural optimization can partially address the problem, but not completely.
3) Increasing the rocket body diameter, increasing the propellant, and reducing the slenderness ratio simultaneously address both the low payload capacity and high maximum dynamic pressure.
3.3 Option C (Three-stage configuration, first stage diameter 3.35m)
1) The first stage adopts a 3.35m diameter and an additional 15 tons of propellant; the roll control system uses a 4x1000N hydrogen peroxide/kerosene auxiliary power system.
2) The second stage is basically the same as Option B, with an additional 2 tons of propellant.
3) The third stage is the same as Option Bo.
4) The fairing diameter is 2.25m or 2.6m and is hoisted integrally. 5) The rocket's total length is 29.237 meters.
6) The takeoff mass is 103.217 tons, the takeoff thrust-to-weight ratio is 1.20, the maximum dynamic pressure is 22.3 kPa, and the maximum axial overload is 5.0 g.
7) The 700 km SSO payload capacity is approximately 1,080 kg, and while meeting domestic measurement and control requirements, the payload capacity is 500 kg.
The main advantages of Option C are:
1) The increased diameter and reduced height result in a slenderness ratio of 8.7, resolving the issue of a high slenderness ratio. This eliminates the limitation of rocket height for dual satellite launches.
2) The increased takeoff mass reduces the maximum dynamic pressure by 39%.
3) The 700 km SSO payload capacity is increased by 25%, and the per-launch cost is comparable to Option B, reducing the unit payload insertion cost. 4) The total length of the first, second, and third stages is 24.2 meters, with a dry weight of 9.02 tons. It can be assembled in the final assembly workshop and shipped directly from the launch station for erection, testing, and fueling before launch, thus enabling rapid launch of small satellites.
5) It can be adapted to existing launch pads. The main issue with Option C is that, compared to the new generation 3.35m module, since the first stage uses a single engine, the overall interface relationship needs to be re-coordinated.
3.4 Option D (Three-stage configuration, first stage diameter 3.0m)
1) The first stage adopts a 3.0m diameter;
2) The second stage is the same as Option C;
3) The third stage is the same as Option C;
4) The fairing diameter is 2.25m or 2.6m, and it is hoisted integrally;
5) The rocket has an overall length of 31.53m and an aspect ratio of 10.51;
6) The takeoff mass is 103.167t, the takeoff thrust-to-weight ratio is 1.202, the maximum dynamic pressure is 22.0kPa, and the maximum axial overload is 5.0g;
7) The 700km SSO payload capacity is approximately 1100kg.
This option has no advantages over Option C.
3.5 Option E (Three-stage Configuration, 3.0m Diameter)
1) The first, second, and third stages have a 3.0m diameter and the same propellant workload as Option C.
2) The second-stage kerosene tank uses a suspended tank with a diameter of 2.25m.
3) The fairing has a diameter of 3.0m and can be hoisted in one piece or in separate sections.
4) The rocket has a total length of 31.19m and a slenderness ratio of 10.4.
5) The takeoff mass is 103.677t, the takeoff thrust-to-weight ratio is 1.196, the maximum dynamic pressure is 21.8kPa, and the maximum axial overload is 4.9g.
6) The 700km SSO payload capacity is approximately 780kg.
Compared to Option D, this option has the following disadvantages: the increased structural mass due to the larger second and third stage diameters reduces the payload capacity. The use of a suspended tank for the second-stage kerosene tank results in an excessively long interstage section, increasing separation difficulty.
3.6 Option F (Three-stage Configuration, 2.5m Diameter)
1) The first, second, and third stages have a 2.5m diameter;
2) The first-stage propellant mass is 71t, with the second and third-stage propellant mass being the same as Option C;
3) The fairing has a diameter of 2.5m and is hoisted integrally;
4) The rocket has an overall length of 34.46m and an aspect ratio of 13.8;
5) The takeoff mass is 98.105t, the takeoff thrust-to-weight ratio is 1.264, the maximum dynamic pressure is 25.8kPa, and the maximum axial overload is 5.0g;
7) The 700km SSO payload capacity is approximately 910kg.
The 2.5m diameter of this option was determined after optimization and can replace the 2.25m module in the original series as a booster for 5m and 3.35m diameter rockets. The three modules in the new generation series have been changed to: 5m, 3.35m, and 2.5m. However, the new diameter module brings new production and process facility problems.
