For manned missions, if conventional chemical propulsion technology is adopted, the initial scale of the earth will reach 1400t, and after nuclear thermal propulsion technology is adopted, the initial scale of the earth can be reduced to 800t t. Nuclear thermal propulsion technology has incomparable advantages in deep space exploration because of its unique performance of high specific impulse and large thrust.
Previous Mars exploration missions showed that there are some necessary conditions for the existence of life on Mars, especially the discovery of water, which greatly stimulated the enthusiasm of human beings to find life on Mars and became a hot spot in international deep space exploration in recent years. Nuclear thermal propulsion technology has incomparable advantages over chemical propulsion technology in deep space exploration because of its unique performance of high specific impulse and large thrust. Moreover, with the gradual development of nuclear power technology, nuclear safety problems can be solved reliably. In order to ensure that China can play a greater role in the field of deep space exploration in the future, it is of great significance to develop nuclear thermal propulsion technology.
In this paper, the overall scheme of nuclear thermal propulsion aircraft is studied under the background of manned boarding mission, and the overall performance, design characteristics and key technologies of nuclear thermal propulsion aircraft are analyzed and sorted out.
With more and more understanding of Mars, NASA, the Russian Federal Space Agency and the European Space Agency have all started scientific research on emigrating to Mars, which is expected to realize the dream of landing on Mars in the mid-1930s. Among them, NASA began the research on manned Mars exploration as early as 1988, and formed the "Mars Reference Mission" (DRM) series of manned landing on Mars.
The design reference system of American manned Mars Exploration 5.0 (MARSRA 5.0) basically established the overall scheme of "heavy launch vehicle+upper stage of nuclear power". The basic scheme is to use seven heavy rockets to send the nuclear thermal propulsion stage and manned/cargo payload into low-earth orbit, and then use two cargo rockets and 1 manned rocket to dock in low-earth orbit respectively, and then be transported to Mars by nuclear thermal propulsion to return to Earth. In the early American manned Mars exploration plan, it was mentioned that the traditional chemical propulsion system was used for manned boarding, and the earth departure scale was as high as1400 t. The structure of nuclear thermal propulsion system was similar to that of chemical rocket engine, and the thrust was almost the same, but the specific impulse was increased to about 900 950s, and the earth departure scale was reduced to 800 t. The principle of "separate transportation of people and goods" was generally adopted in Mastra 5.0 scheme.
Marsra 5.0 manned fire-proof boarding program in the United States
Referring to the MARDRA 5.0 scheme of the United States, China has also made a preliminary manned mission planning. Considering that the initial scale of the earth is 700 800t, * * * will carry out 7 8 launches and 5 docking in low earth orbit.
1) A heavy-duty launch vehicle 1 put the nuclear thermal propulsion and fire-running orbital transfer stage 1 into low-earth orbit;
2) The heavy launch vehicle 2 will push nuclear heat into low earth orbit;
3) Orbital module 1 (Mars lander and elevator) is sent into low earth orbit by heavy carrier rocket 3;
4) Orbital module 2 (life module on the surface of Mars and rover) is sent into low earth orbit by heavy carrier rocket 4;
5) A heavy-duty launch vehicle 5 is used to send the nuclear thermal propulsion into a low-earth orbit in three stages;
6) sending the liquid hydrogen tank into low-earth orbit by a heavy-duty carrier rocket 6;
7) The manned ferry spacecraft (including the spacecraft 2) is sent into the low-earth orbit by the heavy carrier rocket 7;
8) Manned spacecraft 1 Launched into low earth orbit by manned rocket.
Orbital module 1 is docked with orbital module 1 in low-earth orbit, and the orbital module 1 is put into orbit by nuclear thermal propulsion orbital module 1, separated from orbital module 1, and then braked by orbital module 1. Docking the nuclear thermal propulsion misfire and orbit module 2 in the near-earth orbit, sending the orbit module 2 into the misfire orbit by the nuclear thermal propulsion misfire and orbit module 2 separated from the misfire and orbit module 2, braking and pneumatically decelerating the orbit module 2, sending the life module and the rover on the surface of Mars into the ring fire orbit and waiting for the manned spacecraft to enter the orbit; The third stage of thermal propulsion, the liquid hydrogen tank, the manned space shuttle and the manned spacecraft 1 are connected to the near-earth orbit in turn, and the astronauts will enter the space shuttle from the manned spacecraft. The third stage of nuclear thermal fire extinguishing (and the liquid hydrogen tank) will send the manned space shuttle and the manned spacecraft into the fire-running orbit and the ring-fire orbit. The manned space shuttle that was put into orbit first docked with the living quarters on the surface of Mars, and the living quarters were separated from other parts of the space shuttle, and then the living quarters and spacecraft 2 landed on the surface of Mars.
After completing the mission, the astronauts entered the orbit of Mars through the upgrade on Mars and Spacecraft 2, and rendezvous and docked with the manned ferry spacecraft and other parts of the manned spacecraft 1. Before returning to the earth, the astronauts entered the manned spacecraft 1, separated from the ferry spacecraft and re-entered the earth directly.
The nuclear thermal propulsion power system is mainly composed of a nuclear heat engine and a pressurized conveying system. At present, the domestic nuclear heat engine is still in the conceptual design stage. The nuclear heat engine is similar to the expander cycle engine using liquid hydrogen in principle, except that the hydrogen-oxygen combustion chamber is replaced by a nuclear reactor. Liquid hydrogen propellant comes out of the storage tank and is pressurized by the pump. It first enters the cooling thrust chamber of the engine cooling water jacket and then gasifies, and then it is divided into two paths: one path directly enters the thrust chamber, and the other path enters the thrust chamber after blowing the turbine. After the hydrogen entering the thrust chamber is heated by the nuclear reactor, it becomes high-temperature and high-pressure gas, which is ejected at high speed through the nozzle to form thrust.
Schematic diagram of nuclear heat engine concept
Specific impulse of (1) nuclear heat engine
The specific impulse of engine is directly proportional to the square root of propellant temperature and inversely proportional to the square root of molecular weight. Due to the limitation of materials and heat transfer, the temperature of combustion chamber generally does not exceed 3000 4000K, so reducing molecular weight is an effective way to improve specific impulse.
The molecular weight of chemical combustion products generally exceeds 10, and the nuclear heat engine can directly heat the low molecular weight medium to high temperature, thus producing high specific impulse. At present, the best working medium of nuclear heat engine is liquid hydrogen, which has good cooling and expansion ability and is a single substance with the smallest molecular weight. In order to maximize the temperature of the medium, the technical level of nuclear fuel rods plays a decisive role in the specific impulse performance. It is the core key technology of nuclear heat engine, and it is also a technology with a big gap between China and foreign countries in the field of nuclear heat engine.
At present, Russia is at the highest level in this field, and its ternary carbide technology can heat hydrogen to more than 2800K, thus achieving engine specific impulse above 900s s. When the engine area ratio is 300 and the nozzle efficiency is 0.96, the specific impulse changes with the increase of hydrogen heating temperature.
(2) Thrust-weight ratio of nuclear heat engine
Due to the existence of nuclear reactor and related shielding layer, the thrust-to-weight ratio of nuclear heat engine is lower than that of conventional liquid rocket engine, but much higher than that of electric propulsion engine. The design value of thrust-to-weight ratio of American nuclear heat engines is as high as 4.8, generally between 3 and 4. The thrust-to-weight ratio of a nuclear heat engine depends on the nuclear-related components, such as reactor, reflector, shielding layer and control mechanism, while the components related to conventional cryogenic engines, such as thrust chamber, nozzle and turbopump, only account for about 10%.
For nuclear heat engine reactor, its components are mainly composed of core (including fuel and moderator, etc.). ), reflector, reactivity control system, shielding and other internal components.
Taking the nuclear heat engine reactor used for manned landing on Mars in the United States as an example, it is estimated that the total mass of the nuclear reactor is about 3422kg, while the engine thrust is about11.2kn, and the thrust-to-weight ratio is 3.3 14. Considering the engine nozzle, turbopump and propellant delivery pipe, the thrust-to-weight ratio of nuclear heat engine in practical engineering application is about 3.
(3) Start-up and shut-down performance of nuclear heat engine
The energy of conventional rocket engine comes from the chemical reaction of propellant, and the process of accelerating accumulation and decelerating release is directly related to the supply of propellant, so it can be started and shut down quickly.
The nuclear heat engine uses the nuclear reactor as its energy source, and its start-up and shutdown process largely depends on the working requirements and characteristics of the reactor, especially during the shutdown of the nuclear reactor, the radiation effect of some products will last for a long time and need to be continuously cooled.
By analyzing the development experience of American nuclear thermal engine, it is found that the start-up and shutdown process of nuclear thermal rocket engine is different from that of conventional rocket engine, especially after the engine is shut down, it needs to maintain a long cold shutdown process.
The start-up and shut-down characteristics of 34-ton nuclear heat engine for lunar ferry are analyzed. The engine is based on the NRX series engine developed by the American "Nuclear Engine for Launch Vehicles" (NERVA). The design total temperature is 236 1k, the design room pressure is 3. 1MPa, the vacuum specific impulse is 822s, and the flow rate under the design thrust is 41.7 kg/s.
1) Start the process. The start-up process of nuclear thermal rocket engine is somewhat similar to that of conventional cryogenic rocket engine, but it takes much longer.
In the first stage of start-up, liquid hydrogen flows through turbopump, thrust chamber, reactor, etc. Under the action of tank pressure, the reactor is in a low power state. This process takes about 25s, and its main function is to fully precool the engine and preheat the reactor.
In the second stage, the engine began to accelerate, the temperature reached the rated working condition, the thrust reached 60% of the rated thrust, and the duration was about 22.7 s;
In the third stage, under the condition of constant total temperature, the pressure of the combustion chamber rises to the rated working condition, the thrust reaches 100%, and the duration is about 3.6s Generally speaking, the starting process of the nuclear heat engine lasts about 52s, and after deducting the pre-cooling time of the engine, it is about 27s. The average specific impulse during start-up is only about 600s s.
2) shutdown process. The shutdown process of nuclear heat engine is basically the reverse process of startup process, but it takes longer. First, the engine power should be reduced to 60%. During this process, the total temperature of the engine remains unchanged, and the pressure of the combustion chamber drops for about 3.6s, and the specific impulse of the engine remains unchanged during this process; Then, the engine is maintained at 1. 3min in this state, the main purpose is to reduce the waste heat generated in the subsequent cold shutdown process, thus saving propellant consumption. Then the total temperature and thrust of the engine will continue to drop until the engine is turned off, and it is necessary to maintain a long-term and low-flow cooling waste heat discharge stage. The whole shutdown process of the 34-ton nuclear heat engine lasted about 350s, and the average specific impulse of the engine was about 600s during the whole shutdown process.
The biggest difference between nuclear heat engine and conventional engine is that there is still a stage of waste heat emission after engine shutdown, mainly because some reaction products are still highly radioactive after reactor shutdown, which will release waste heat. Take the nuclear heat engine of a 34-ton lunar ferry as an example. The process lasts about 64 hours, the thrust is about 134N, and the specific impulse is about 400s s. Due to the long duration, the consumption of liquid hydrogen needs to be considered. At the same time, the cooled hydrogen in this process can be designed to generate electricity and provide a certain power source for the whole machine.
The nuclear reactor will emit gamma rays and a large number of neutrons during its operation, which will do harm to the electronic components and astronauts on the spacecraft, so it needs to be shielded to reduce its radiation level to below the allowable value. For reactors used in space, electronic components and astronauts are in a relatively concentrated position due to the strict limitation of volume and mass, so shadow shielding can be used to keep the radiation level at a low level.
For spacecraft using nuclear power, it is generally designed as a slender structure, that is, the instrument cabin and personnel cabin are located at one end, the nuclear reactor is located at the other end, and the liquid hydrogen storage tank is located between the two ends.
Because the linear motions of neutrons and gamma rays are specific and the positions that need shielding are relatively concentrated, it is necessary to place the shielding area in the shadow area of the shielding block.
Schematic diagram of radiation shielding arrangement
According to the protection indexes formulated during the overhaul of Daya Bay and Qinshan nuclear power plants, the collective dose does not exceed 600 (msv), and the maximum individual dose does not exceed 15mSv. Considering that the last stage of nuclear thermal propulsion is limited by volume and mass, its radiation level may be slightly higher. Assuming that the allowable leakage value of the radiation safety zone of the nuclear thermal propulsion system is less than 20mSv per day, this value has greatly exceeded the radiation protection index requirements formulated in the overhaul of Daya Bay and Qinshan nuclear power plants.
Considering that the mission period of Mars exploration is three years, and assuming that all the above radiation is absorbed by rocket electrical products, the cumulative absorbed dose during the whole mission period is 2 1.9j/kg. Under the current product level, the dose of ionizing radiation that non-radiation-resistant semiconductor components can bear is not less than 100J/kg.
It can be seen that the radiation dose of rocket electrical products is less than the bearing capacity of components, and the nuclear thermal propulsion has no essential influence on the electrical system scheme, but the nuclear thermal engine must have the basic radiation shielding ability to control the external radiation within an acceptable range.
For deep space exploration mission, the complex deep space radiation environment is the main environment faced by spacecraft, and the deep space environment exposed outside the geomagnetic layer is full of high-energy mixed space radiation.
Layout of nuclear thermal propulsion spacecraft
According to the flight phase of spacecraft in deep space, the deep space environment can be divided into three parts:
First, the space radiation environment during the flight from the earth to other planets, the main radiation sources are solar particle events and galactic cosmic rays;
The second is the space radiation environment of spacecraft during the landing of stars. The main radiation sources are solar cosmic rays and galactic cosmic ray particles captured by the magnetic field of stars.
The third is the radiation environment on the surface of the star where the spacecraft landed, mainly the secondary radiation after the star absorbed cosmic radiation.
The hazards caused by deep space radiation environment are mainly radiation damage and single event. The interaction between high-energy electrons, protons and a small amount of heavy ions in deep space radiation environment will cause damage and destruction to spacecraft materials, in which high-energy electrons ionize spacecraft materials, high-energy protons and heavy ions ionize and replace spacecraft materials.
When designing the electrical system of deep space exploration spacecraft, the calculation error caused by single event caused by photothermal radiation or the risk of changing the value in memory should be considered. This situation should be considered in software design, and methods such as calculation redundancy and error checking should be used to detect and judge to ensure the correctness of rocket engine calculation.
The working environment of the upper stage of nuclear thermal propulsion is outside the atmosphere and will not be affected by aerodynamic load, so its structural scheme design can not be limited by aerodynamic shape. Taking the concept diagram of nuclear thermal power carrier released by Russia as an example, the main load-bearing structure of the carrier is mainly rod system to improve the structural efficiency of the carrier. Moreover, because there is no space limitation of the fairing, the structure of the payload is more flexible and there are more spatial distribution schemes.
Nuclear thermal propulsion system only needs liquid hydrogen as working medium, so it only needs liquid hydrogen as storage tank, and there is no need to set oxidant storage tank, which has less constraints on structural design and can better optimize structural scheme.
However, after the nuclear heat engine is used, it will bear worse high-temperature environmental conditions than the conventional engine, so it is necessary to fully consider the thermal protection requirements of structures, instruments and cables near the engine in the structural design process to ensure the normal work of each system and single machine.
Moreover, compared with the conventional engine, the structure of the nuclear heat engine is larger, and it is necessary to increase the structural strength of the engine part, especially around the reactor, to ensure the tightness of each part of the engine.
Concept map of Russian nuclear thermal power vehicle
Referring to the Mars DRA5.0 scheme of the United States, a preliminary manned boarding scheme similar to that of the United States is proposed. The total starting scale of the earth is about 700 ~ 800 tons, and the fire transfer is completed in three times, and the single starting scale of the earth is about 300 tons. By analyzing the launch efficiency, working time, gravitational loss and orbital quality when the energy C3e from parking orbit to the Earth is 8 or 20km2/s', respectively, the thrust scale of the last stage of nuclear thermal propulsion and the overall parameter suggestions of the nuclear heat engine are given.
Assuming that the entry orbit is a near-earth circular orbit with a height of 200km, the thrust-to-weight ratio of the nuclear engine and the thermal engine is 3, and the specific impulse is 905s s. Considering the influence of gravity loss, the launch efficiency of the nuclear thermal propulsion vehicle at different thrust scales is analyzed, in which the launch efficiency refers to the ratio of the on-orbit mass (after deducting the dry weight of the nuclear engine, it enters the ground fire transfer orbit) to the initial mass of the entry orbit. It can be seen that the emission efficiency is the highest when the overload is between 0.13 and 0.16.
When different overload is considered in the launch efficiency, different orbital transfer time will bring the influence of gravitational loss. The specific influence is that the smaller the overload, the longer the working time, the greater the gravity loss, but the smaller the dry weight of the engine. According to the starting scale of a single fire transfer of 300t, the thrust of nuclear thermal propellant launch vehicle should be about 45 t. Combined with the research situation of nuclear thermal engines in the United States and Russia, it is suggested that the thrust of nuclear thermal engine should be 15t, and the nuclear thermal propulsion launch vehicle should be connected in parallel with three engines.
Variation of Earth Transfer Emission Efficiency with Overload
Nuclear thermal propulsion technology has incomparable advantages in the future deep space exploration mission because of its characteristics of large thrust and high specific impulse, but it should also be noted that the engineering application of nuclear thermal technology still has a long way to go and many technical problems need to be overcome. According to the analysis of manned missions based on nuclear thermal propulsion, it takes about 180 days for a nuclear thermal propulsion vehicle to reach Mars from Earth. After staying on Mars for a period of time (ranging from one week to one and a half years), the nuclear heat engine is re-ignited and returned to Earth, so the long-term storage time of propellant should be at least half a year, which is a great challenge to the existing long-term storage technology of liquid hydrogen.
In addition, the specific heat of the high-temperature gas thrust by the nuclear heat engine (about 20 000 kJ/kg k at a total temperature of 2500K) is much higher than that of the traditional hydrogen-oxygen engine (about 3 400 kJ/kg k at a total temperature of 3400K), which leads to a higher wall heat flux than that of the traditional engine and brings great difficulties to cooling.
Therefore, in order to realize the application of nuclear thermal propulsion in manned boarding mission, it is necessary to focus on solving major technical problems such as miniaturization of nuclear thermal reactor, cooling of thrust chamber of nuclear thermal engine and long-term storage of propellant.