A Preliminary Study on the Overall Scheme of Nuclear Thermal Propulsion Upper Stage for the Manned Mars Launch Vehicle

For the manned mission to the fire, if the conventional chemical propulsion technology is used, the size of the earth departure reaches 1400t, while after adopting nuclear thermal propulsion technology, the earth departure scale can be reduced to 800t. Nuclear thermal propulsion technology, with its unique performance of high specific impulse and large thrust, has the advantage of deep space exploration that chemical propulsion rockets can not be compared.

Previous Mars exploration missions have shown that Mars has some of the necessary conditions for the existence of life, especially the discovery of water, which has greatly stimulated the enthusiasm of human beings to search for life on Mars, and has become a hotspot for international deep space exploration in recent years. Nuclear thermal propulsion technology, with its unique performance of high specific impulse and large thrust, has the advantages of deep space exploration that are incomparable to chemical propulsion technology. Moreover, with the gradual development of nuclear power technology, the safety of nuclear energy can be reliably solved. In order to ensure that our country can play a greater role in the field of deep space exploration in the future, the development of nuclear thermal propulsion technology is of great significance.

This paper takes the manned mission to the fire as the background, and conducts a preliminary study on the overall program of the nuclear thermal propulsion vehicle, and carries out a preliminary analysis and sorting out of the overall performance, design characteristics, and key technologies of the nuclear thermal propulsion vehicle.

As human beings know more and more about Mars, NASA, the Russian Federal Space Agency, the European Space Agency have begun to migrate to Mars scientific research, is expected to realize the dream of human beings landing on Mars in the middle of the 21st century 30s. Among them, the National Aeronautics and Space Administration as early as 1988 has begun a manned Mars exploration program research, and the formation of a manned landing on Mars "Mars reference mission" (DRM) series of programs.

The U.S. "manned Mars Exploration Design Reference System 5.0" (Mars DRA5.0), basically established a "heavy-lift launch vehicle + nuclear-powered upper stage" of the overall program, the basic program for the use of 7 heavy rockets will be a nuclear thermal propulsion stage, manned/cargo payloads to the near-Earth orbit, and then docked into two cargo rockets respectively, the near-Earth orbit. The basic program is to use seven heavy rockets to send the nuclear thermal propulsion stage and the manned/cargo payload to the near-Earth orbit, and then dock them into two cargo rockets and one manned rocket in the near-Earth orbit, which will be transported to Mars by the nuclear thermal propulsion and return to Earth. Early, the U.S. manned Mars exploration program has mentioned the use of traditional chemical propulsion system for manned boarding of the fire, the earth departure scale as high as 1400t. Nuclear thermal propulsion system structure and chemical rocket engine is similar to the thrust is roughly the same, but the ratio of impulse increased to 900 950s or so, the earth departure scale can be reduced to 800t. Mars DRA5.0 program overall to take the "manned transportation, cargo, and cargo payloads, and return to earth". The overall program of Mars DRA5.0 adopts the principle of "separate transportation of people and cargo, the first thing after people".

U.S. Mars DRA5.0 manned boarding fire program

Reference to the U.S. Mars DRA5.0 program, our country has also carried out a preliminary manned boarding fire mission planning, according to the earth departure scale of 700 800t considerations, *** for 7 8 launches, in near-Earth orbit for 5 docking.

1) by the heavy launch vehicle 1 will be nuclear thermal propulsion Pentium orbit stage 1 into near-Earth orbit;

2) by the heavy launch vehicle 2 will be nuclear thermal propulsion Pentium orbit stage 2 into near-Earth orbit;

3) by the heavy launch vehicle 3 will be the orbital module 1 (Mars Landing Descender and Ascender) into near-Earth orbit;

4) by the heavy launch vehicle 4 will be the Orbital Module 2 (Mars Surface Living Module and Mars Rover) into near-Earth orbit (NEO) by Heavy Launch Vehicle 4;

5) Nuclear Thermal Propulsion (NTP) Penfold Orbital Stage (POS) 3 into NEO by Heavy Launch Vehicle 5;

6) Liquid Hydrogen Tank (LIHT) into NEO by Heavy Launch Vehicle 6;

7) Crewed Swinging Spacecraft (CST) (with Spaceship 2) into NEO by Heavy Launch Vehicle 7;

8) the manned spacecraft 1 into near-Earth orbit by a manned rocket.

Docking of nuclear thermal propulsion Penfire orbital stage 1 and orbital module 1 in near-Earth orbit, sending orbital module 1 into Penfire orbit by nuclear thermal propulsion Penfire orbital stage 1, separation of orbital module 1 from Penfire orbital stage 1, and after that, the descender and the ascender will be sent into circumferential orbit by the braking and aerodynamic deceleration of the orbital module 1, and the descender and the ascender will be landed on the surface of the Mars; docking of nuclear thermal propulsion Penfire orbital stage 2 and orbital module 2 in near-Earth orbit, and sending manned ferry spacecraft 1 (including spacecraft 2) into near-Earth orbit by manned rocket. Docking of nuclear thermal propulsion Pentium stage 2 and orbital module 2 in near-Earth orbit, sending orbital module 2 into Pentium orbit by nuclear thermal propulsion Pentium stage 2, separating orbital module 2 from Pentium stage 2, and after that, the Mars surface living module and rover will be sent into circumferential orbit by the braking and pneumatic deceleration of orbital module 2, waiting for the subsequent orbiting manned spacecrafts; Docking of thermal propulsion Pentium stage 3, liquid hydrogen tank, manned ferry spacecraft and manned spacecraft 1 in near-Earth orbit in sequence, and the astronauts will enter into the ferry spacecraft from the manned spacecrafts. The manned spacecraft enters the ferry vehicle, and the nuclear thermal Penfire stage 3 (and the liquid hydrogen storage tank) sends the manned ferry spacecraft and the manned spacecraft into the Penfire orbit and the Ring of Fire orbit. The manned ferry vehicle and the orbiting Mars Surface Life Module (MSLM) docked in Rimfire orbit, the MSLM separated from the rest of the ferry vehicle, and then the MSLM and Spacecraft 2 landed on the surface of Mars.

After completing their mission, the astronauts entered Mars orbit via Mars Upgrade and Spacecraft 2, and rendezvoused and docked with the rest of the Manned Ferry Vehicle and Manned Spacecraft 1. Before returning to Earth, the astronauts enter Manned Spacecraft 1, separate from the ferry spacecraft, and re-enter Earth directly.

The nuclear thermal propulsion power system mainly consists of two parts: the nuclear thermal engine and the pressurized delivery system. At present, the domestic nuclear thermal engine is still in the conceptual design stage, the nuclear thermal engine in principle with liquid hydrogen as the mass of the expansion cycle engine is similar, the difference is that the hydrogen-oxygen combustion chamber replaced by a nuclear reactor. Liquid hydrogen propellant from the storage tank out of the pump pressurization first into the engine cooling jacket cooling thrust chamber after gasification, and then divided into two ways: one way directly into the thrust chamber, the other way to blow the turbine into the thrust chamber. Into the thrust chamber of the hydrogen heated by the nuclear reactor, into a high-temperature, high-pressure gas through the nozzle high-speed spray, the formation of thrust.

Conceptual schematic of a nuclear thermal engine

(1) Nuclear thermal engine specific impulse

The engine specific impulse is proportional to the propulsion medium temperature of the open square, inversely proportional to the molecular weight of the open square. Due to material and heat transfer limitations, the combustion chamber temperature generally does not exceed 3000 4000K, so reducing the molecular weight is an effective way to improve the specific impulse.

The molecular weight of chemical combustion products generally exceeds 10, while nuclear thermal engines can directly heat low molecular weight media to high temperatures, resulting in high specific impulse. Currently, the best working medium for nuclear thermal engines is liquid hydrogen, which has both good cooling and expansion capabilities and is the smallest molecular weight monomer. In order to maximize the temperature of the medium, the level of nuclear fuel rod technology plays a decisive role in comparison with the impulse performance, is the core of the nuclear heat engine is the most critical technology, but also in the field of China's nuclear heat engine and the gap between foreign countries in the larger technology.

At present, Russia is at the highest level in this field, its ternary carbide technology can be heated to more than 2800K hydrogen, so as to realize the engine specific impulse more than 900 s. In the case of the engine area ratio of 300 and the nozzle efficiency of 0.96, with the increase in the hydrogen heating temperature, the specific impulse changes accordingly.

(2) nuclear thermal engine thrust mass ratio

Nuclear thermal engine due to the existence of nuclear reactors and related shielding layer, thrust mass ratio is lower than conventional liquid rocket engines, but much larger than the electric propulsion engine, the United States nuclear thermal engine thrust mass ratio design value of up to 4.8, generally taken in 3 4 between. Nuclear thermal engine thrust mass ratio depends on nuclear-related components, such as reactors, reflectors, shielding, control mechanisms, etc., and conventional cryogenic engine-related components, such as the thrust chamber, nozzle, turbopumps and other mass accounts for only about 10%.

For the reactor of a nuclear thermal engine, the components are mainly composed of the core (with fuel and moderator, etc.), the reflector layer, the reactivity control system, the shielding, and other in-heap components.

Take the U.S. manned landing on Mars with a nuclear thermal engine reactor as an example, by estimation, the total mass of the nuclear reactor is about 3422kg, while the engine thrust is about 111.2kN, the thrust-to-mass ratio is 3.314. Considering the engine nozzle, turbopumps, and the propellant delivery tube, etc., the actual engineering application of the nuclear thermal engine thrust-to-mass ratio of 3.

(3) nuclear thermal engine startup and shutdown performance

Conventional rocket engines derive their energy from the chemical reaction of the propellant, and their accelerated accumulation and decelerated release process is directly related to the supply of propellant, so they can realize relatively fast startup and shutdown.

The nuclear thermal engine uses a nuclear reactor as the energy source, and its startup and shutdown process depends largely on the reactor's operating requirements and characteristics, especially during the shutdown process of the nuclear reactor, some of the products of the radiation effect will continue for a long time, and need to be cooled continuously.

By analyzing the U.S. experience in the development of nuclear thermal engines, the startup and shutdown process of nuclear thermal rocket motors is somewhat different from that of conventional rocket motors, especially after the engine shutdown, but also to maintain a longer period of cold shutdown process.

A preliminary analysis of the start-up and shutdown characteristics of the 34-ton lunar ferry nuclear thermal engine, which is based on the U.S. "Nuclear Engine for Launch Vehicles" (NERVA) program development of the NRX series of engines as a prototype, the design of the total temperature of 2361K, the design of the chamber pressure of 3.1MPa, vacuum, specific impulse of 822s, the design of the thrust of 822s, the specific impulse of 822s, the design of the thrust of the rocket engine. Specific impulse 822s, the flow rate under the designed thrust is 41.7kg/s.

1) Starting process. Nuclear thermal rocket motor starting process and conventional low-temperature rocket motor is somewhat similar, but much longer.

The first stage of startup, liquid hydrogen in the reservoir pressure flow through the turbopump, thrust chamber, reactor, etc., the reactor is at a lower power, the process takes about 25s, the main role is to fully pre-cool the engine, and will be the reactor preheat.

The second stage of the engine starts to accelerate the start, the temperature reaches the rated condition, the thrust reaches 60% of the rated thrust, which lasts about 22.7s;

The third stage is in the case of the total temperature remains unchanged, the chamber pressure increases to the rated condition, and the thrust reaches 100%, which lasts about 3.6s.

Overall, the start-up process of the nuclear thermal engine lasts about 52s, minus the engine After deducting the engine pre-cooling time, it takes about 27s, and the average specific impulse of the starting process is only about 600s.

2) Shutdown process. The shutdown process of a nuclear thermal engine is basically the inverse of the starting process, but it takes a little longer. First, the engine is reduced to 60% power. This process of total engine temperature remains unchanged, the chamber pressure is reduced, lasted about 3.6s, the process of the engine specific impulse is unchanged; and then, the engine in this state to maintain 1 3min, the main purpose is to reduce the amount of waste heat generated in the subsequent cold shutdown process, in order to save the propellant consumption; and then, the total engine temperature, thrust and then continue to decline to the engine shutdown, but also need to maintain a long period of time to maintain a small flow of cooling of the waste heat emission phase. The entire shutdown process of the 34-ton nuclear thermal engine lasted about 350 s. The average specific impulse of the engine during the entire shutdown process was about 600 s.

The biggest difference between the nuclear thermal engine and the conventional engine lies in the fact that there is still a waste heat emission phase after the engine shutdown, which is mainly due to the fact that after the reactor shutdown, some of the reaction products are still highly radioactive and will release waste heat. Take the 34-ton lunar swing with nuclear thermal engine as an example, the process lasts about 64h, the thrust is about 134N, the specific impulse is about 400s, because of the long duration, this process of liquid hydrogen consumption needs to be considered, at the same time, the cooling hydrogen of this process can be designed for power generation to provide a certain source of electric power for the entire vehicle.

The nuclear reactor will emit γ-rays and a large number of neutrons during operation, which will be hazardous to the electronic components on the spacecraft and the astronauts, and therefore need to be shielded to reduce the radiation level to below the permissible value. For reactors used in space applications, where electronic components and astronauts are in a relatively centralized location due to tighter volume and mass limitations, shadow shielding can be used to keep radiation levels low.

For spacecraft using nuclear power, they are generally designed as elongated structures, i.e., the instrumentation module and crew module are located at one end, the nuclear reactor is located at the other end, and the liquid hydrogen storage tank is between the two ends.

Due to the specific linear motion of neutrons and γ-rays, and the relative concentration of locations to be shielded, it is necessary to place the shielded area in the shaded area of the shielding block.

Schematic diagram of radiation shielding arrangement

Reference to Daya Bay and Qinshan nuclear power plant overhaul of the development of the protection index, the collective dose does not exceed 600 (people - mSv), the maximum dose of an individual does not exceed 15mSv, taking into account the nuclear thermal propulsion end stage by the volume of the mass of the limitations, the radiation level may be slightly higher, assuming that the nuclear thermal propulsion system radiation safety zone of the permissible leakage value of less than daily 20mSv, this value has greatly exceeded the requirements of the radiation protection indexes developed during the overhaul of the Daya Bay and Qinshan nuclear power plants.

In accordance with the Mars exploration mission cycle for three years, and assuming that the above radiation is fully absorbed by the rocket electrical products, the cumulative absorbed dose for the entire mission cycle is 21.9J/kg, in the current level of products, non-radiation-resistant semiconductor components can withstand the dose of ionizing radiation is not less than 100J/kg.

It can be seen that the rocket electrical products are subjected to radiation dose is less than the components of the ability to withstand, nuclear thermal propulsion on the electrical system program does not have an intrinsic impact, but the nuclear thermal engine must have the basic radiation shielding capabilities, the external radiation control to an acceptable range.

For deep space exploration missions, the complex deep space radiation environment is the main environment faced by spacecraft, and deep space environments exposed beyond the geomagnetosphere are filled with high-energy mixed space radiation.

Arrangement of a spacecraft using nuclear thermal propulsion

The deep space environment can be divided into three parts according to the flight phase of the spacecraft in deep space:

First, the space radiation environment during the journey from the Earth to other planets, whose main sources of radiation are the solar particle events and the galactic cosmic rays;

Second, the space radiation environment during the process of spacecraft landing on the stars, whose main sources of radiation are the magnetic field of the stars;

Second, the space radiation environment during the process of landing on the stars, whose main sources of radiation are the magnetic field of the stars. The main sources of radiation are solar cosmic rays and galactic cosmic ray particles captured by the magnetic field of the star;

The third is the radiation environment on the surface of the star landed on by the spacecraft, which is mainly the secondary radiation occurring after the absorption of cosmic radiation by the star.

The hazards caused by the deep space radiation environment are mainly radiation damage and single-particle events. The deep space radiation environment is full of high-energy electrons, protons, and a small amount of heavy ions and spacecraft materials, which will cause damage and destruction of the performance of spacecraft materials, in which the high-energy electrons produce ionization of spacecraft materials, and the high-energy protons and heavy ions produce ionization and displacement of spacecraft materials.

In the design of the electrical system of the spacecraft for deep space exploration, it is necessary to consider the risk of calculation errors caused by single-particle events due to photothermal radiation, or changing the value in the memory, etc. The software design needs to take this situation into account, and use calculation redundancy, error checking, and other methods of detecting and discriminating, to ensure that the arrows of the machine calculations are correct.

The nuclear thermal propulsion upper stage operates outside the atmosphere and is not subject to aerodynamic loads, so its structural program design can be independent of aerodynamic profile limitations. Taking the conceptual diagram of the nuclear thermopower carrier released by Russia as an example, the main bearing structure of the carrier is based on the rod system, so as to improve the efficiency of the carrier structure. And because there is no fairing space limitations, the payload structure form more flexible, more space distribution programs.

The nuclear thermal propulsion system requires only one type of liquid hydrogen, so only one type of liquid hydrogen tank is needed, and there is no need to set up another oxidizer tank, so there are fewer constraints on the structural design, and the structural optimization of the scheme can be carried out in a better way.

However, with a nuclear thermal engine, the engine will be subjected to harsher high-temperature environmental conditions than a conventional engine, which requires comprehensive consideration of the thermal protection needs of the structures, instruments, and cables near the engine during the structural design process to ensure that the systems and standalone machines work properly.

And compared to conventional engines, the nuclear thermal engine structure is more bulky, which requires an increase in the engine part, especially the structural strength around the reactor, while ensuring the sealing of the engine components.

Conceptual diagram of the Russian nuclear thermal power carrier

Reference to the U.S. Mars DRA5.0 program, proposed with the United States similar to the manned boarding of the fire preliminary program, the Earth's total departure size of about 700 ~ 800t, three times to complete the transfer of the Earth and fire, a single Earth departure size of about 300 tons. By analyzing the launch efficiency, working time, gravitational loss, and in-orbit mass when accelerating from the parking orbit to the Earth departure energy C3e of 8 or 20km2/s' respectively, the thrust scale of the nuclear thermal propulsion upper stage as well as the overall parameters of the nuclear thermal engine are given as suggestions.

Assuming that the mooring orbit is a near-Earth circular orbit with an altitude of 200km, the nuclear . Thermal engine thrust-to-mass ratio is taken as 3, specific impulse is taken as 905s, considering the effect of gravitational loss, the launch efficiency of nuclear thermal propulsion carrier is analyzed in different thrust scale cases, where the launch efficiency refers to the ratio of the orbit entry mass (into the ground fire transfer orbit) to the departure mass of the mooring orbit after deducting the dry weight of the nuclear thermal engine. It can be seen that the launch efficiency is highest when the overload is between 0.13 and 0.16.

In the launch efficiency has taken into account the case of different overloads, the different time of the orbit change brings the impact of gravitational loss, the specific impact is the smaller the overload, the longer the working time, the greater the gravitational loss, but the engine dry weight is smaller. In accordance with the single Earth-fire transfer departure size of 300t considerations, the nuclear thermal propellant carrier thrust should be in the 45t or so the best, combined with the United States, Russia, nuclear thermal engine research, it is recommended that the nuclear thermal engine thrust in accordance with the 15t considerations, the nuclear thermal propellant carrier in accordance with the 3 machines in parallel.

Earth transfer launch efficiency changes with overload

Nuclear thermal propulsion technology with its large thrust, high specific impulse and other characteristics of the future deep space exploration mission has an incomparable advantage, but should also be seen, at present, from the engineering application of nuclear thermal technology there is still a long way to go, there is still a need to overcome a lot of technical difficulties. According to the current analysis of the manned mission based on nuclear thermal propulsion, nuclear thermal propulsion vehicle from the Earth to Mars needs about 180 days, in Mars to stay - - after a period of time (from one week to a year and a half, etc.), the nuclear thermal engine and then ignited to return to the Earth, so that the propellant long term storage should be at least half a year, which is a great challenge for the existing long-term storage technology of liquid hydrogen.

In addition, the nuclear thermal engine thrust high-temperature gas-hydrogen specific heat (total temperature of 2500K when about 20000kJ/kg K) to be much higher than the traditional hydrogen-oxygen engine high-temperature gas specific heat (gas total temperature of 3400K, gas specific heat of 3000kJ/kg K or so), resulting in the wall of the heat flux density is higher than that of traditional engines, which gives the cooling brings great difficulties.

Therefore, to realize the application of nuclear thermal propulsion in the manned boarding fire mission, need to focus on solving the nuclear thermal reactor miniaturization, nuclear thermal engine thrust chamber cooling, long-term storage of propellant and other major technical challenges.