CN108423196A - The two-stage that the first order can be reused is entered the orbit the method for entering the orbit of spacecraft - Google Patents
The two-stage that the first order can be reused is entered the orbit the method for entering the orbit of spacecraft Download PDFInfo
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- MYMOFIZGZYHOMD-UHFFFAOYSA-N Dioxygen Chemical compound O=O MYMOFIZGZYHOMD-UHFFFAOYSA-N 0.000 claims description 10
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- B—PERFORMING OPERATIONS; TRANSPORTING
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Abstract
Enter the orbit the invention discloses the two-stage that a kind of first order can be reused the method for entering the orbit of spacecraft, including step 1, two-stage is entered the orbit the model construction of spacecraft：The first order is sky tower, using SABRE engines as power plant；The second level is pencil rocket, and payload accounts for the 5.89% of gross mass of taking off；Step 2, level-one section of taking off；Step 3, level-one air-breathing mode section rises to 25km with air-breathing mode；Step 4, one-stage rocket schema section, SABRE engines are switched to rocket mode and two-stage spacecraft of entering the orbit are pushed to continue to rise to 100km, the first order and second level separation, and the mode for gliding that the first order is unpowered, which returns to ground, to be realized and reuse；Step 5, the two level section second level of climbing is climbed vertically, and payload is sent into the LEO of 300km with the speed more than 7700m/s.The present invention is based on SABRE engines on the basis of considering technical difficulty and launch cost, realizes that two-stage is entered the orbit, and the first order can be reused；And payload is made to obviously increase.
Description
Technical Field
The invention relates to the technical field of spacecraft orbit entering, in particular to an orbit entering method of a two-stage orbit entering spacecraft, wherein the first stage of the two-stage orbit entering spacecraft can be repeatedly used.
Background
A coordinated air-breathing rocket Engine (SABRE) proposed by Reaction engineering Ltd, british, is a precooling type combined cycle Engine which cools incoming air by using a low-temperature medium, and is formed by organically combining a turbine Engine, a rocket Engine and a ramjet. The combined type air-vehicle power device integrates the advantages of high thrust-weight ratio, wide working range and high specific impulse of a turbine engine of the rocket engine, has two working modes of air suction and the rocket, and is expected to become a novel power device of a single-stage/multi-stage orbit-entering reusable air-vehicle.
The single-stage orbit input can greatly reduce the emission cost, but is difficult to realize, mainly because a large amount of structural mass is brought into the space, and the requirement on the structural coefficient is extremely strict. The structure quality of failure is continuously abandoned in the multi-stage track entering, the realization is simpler, and the cost is increased.
The invention provides a two-stage orbit entering scheme based on a SABRE engine, and provides a power model and a pneumatic model of an aerospace craft; a motion equation is established in stages, numerical simulation research is carried out, technical feasibility is analyzed, and a result can provide reference for developing research on an on-orbit scheme of a precooling-based combined cycle engine in China.
Disclosure of Invention
The invention provides a two-stage orbit entering spacecraft orbit entering method with a reusable first stage aiming at the defects of the prior art, the two-stage orbit entering spacecraft orbit entering method with the reusable first stage realizes two-stage orbit entering based on an SABRE engine on the basis of comprehensively considering technical difficulty and emission cost, and the first stage can be reused; in addition, on the premise that the initial takeoff mass is not changed, the effective load can be obviously increased, and the technical difficulty is low.
In order to solve the technical problems, the invention adopts the technical scheme that:
a method for the first-stage reusable two-stage orbit spacecraft orbit comprises the following steps.
Step 1, constructing a model of a two-stage orbit spacecraft: the two-stage orbit spacecraft comprises a first stage and a second stage; the first-stage configuration is a trumpet creeper, and a SABRE engine is adopted as a power device; the second stage is a small rocket, and a liquid hydrogen/liquid oxygen rocket engine is used as a power device; the payload in the second stage accounts for 5.89% of the total takeoff mass.
Step 2, a first-stage takeoff section: the two-stage in-orbit spacecraft leaves the ground at a fixed elevation angle theta from the vicinity of the equator of the earth, and the flight time of the one-stage takeoff section is t 1.
Step 3, a first-stage air suction mode section: after the primary takeoff section is finished, the aircraft climbs to the height of not less than 25km in an air suction mode under the action of gravity, aerodynamic force and thrust, and the flight time of the primary air suction mode section is t 2.
Step 4, a first-stage rocket mode section: after the first-stage air suction mode section is finished, the SABRE engine is switched to a rocket mode and pushes the two-stage in-orbit spacecraft to continuously climb, after the altitude reaches 100km, the first stage and the second stage are separated, and the first stage returns to the ground in an unpowered gliding mode to realize repeated use; the time of flight of the first stage rocket mode segment is t 3.
Step 5, a secondary climbing section: after the first and second stages are separated, the second stage climbs vertically and makes a gravity turn at the end of the trajectory and sends the payload into a 300km near-earth orbit at a speed exceeding 7700 m/s; the time of flight of the secondary climb segment is t 4.
In the step 3, when the SABRE engine is in the first-stage air suction mode section, the thrust of the SABRE engine is increased along with the rise of the height; and 4, keeping the thrust of the SABRE engine unchanged when the SABRE engine is in a first-stage rocket mode section.
In step 2, the equation of motion of the first-stage takeoff section is expressed as follows:
in the formula: v represents the aircraft speed; t represents the thrust of the engine borne by the aircraft in the first-stage takeoff section; d represents the aerodynamic resistance of the aircraft; f denotes an aircraftThe ground friction resistance; l represents the aerodynamic lift experienced by the aircraft; m represents the aircraft mass; m is0Representing an initial mass of the aircraft;representing an aircraft fuel consumption rate; ρ represents the atmospheric density of the aircraft at the position; s represents the aircraft characteristic area; cDRepresenting an aircraft drag coefficient; cLRepresenting the lift coefficient of the aircraft; f represents a ground friction coefficient; w represents the weight force borne by the aircraft; x represents the aircraft horizontal displacement.
In step 3, the equation of motion of the first-stage inspiration mode segment is expressed as follows:
wherein,
in the formula, T represents the thrust of an engine borne by the aircraft in a first-stage air suction mode section; m is1Representing the initial total mass of the aircraft in the primary air induction mode segment;representing an aircraft fuel consumption rate; g0 represents the earth surface gravitational acceleration; k represents a lift-drag ratio coefficient; l represents the aerodynamic lift experienced by the aircraft; d represents the aerodynamic resistance of the aircraft; cLRepresenting the lift coefficient of the aircraft; cDRepresenting the aircraft drag coefficient.
In step 4, the equation of motion of the first-stage rocket mode section is expressed as follows:
wherein,
in the formula, T represents the engine thrust borne by the aircraft in the first-stage rocket mode section; m is1Representing the initial total mass of the aircraft in a stage one rocket mode segment;representing an aircraft fuel consumption rate; g0R, H respectively representing the acceleration of gravity on the earth's surface, the radius of the earth, and the altitude at which the aircraft is located; k represents a lift-drag ratio coefficient; l represents the aerodynamic lift of the aircraft; d represents the aerodynamic resistance of the aircraft; cLRepresenting the lift coefficient of the aircraft; cDRepresenting an aircraft drag coefficient; theta represents the aircraft angle of attack.
In step 5, the equation of motion of the secondary climbing section is expressed as follows:
in the formula: t represents the thrust of the engine borne by the aircraft in the secondary climbing section; w represents the weight force borne by the aircraft; m is2Representing the initial total mass of the aircraft when the rocket mode climbs to 100 km;represents a fuel consumption flow rate; isp represents the liquid oxygen/liquid hydrogen rocket engine specific impulse.
t 1-46 s, t 2-385 s, t 3-304 s, t 4-33 s, total time of flight 768 s.
In step 5, the second stage climbs vertically and makes a gravity turn at the end of the trajectory and sends the payload into a 300km near-earth orbit at 7925.009 m/s.
In the step 1, a liquid hydrogen/liquid oxygen rocket engine is adopted as a power device in the second stage, and when the height of the rocket engine is higher than the design height of a spray pipe, the specific impulse is kept to be vacuum specific impulse and the numerical value is not changed any more; the vacuum specific impulse of the oxyhydrogen rocket is 4500m/s, and the thrust of the rocket engine is controlled by adjusting the fuel flow in the flight process.
The invention has the following beneficial effects:
(1) when the aircraft reaches the track height of 300km, the speed is 7909.964m/s, the requirement of rail entering speed is met, and two-stage rail entering can be realized.
(2) The aircraft gross takeoff weight 345t and payload 20.305 t. The first grade is 289.98t, and the structural coefficient is 0.1902. The second level weighs 55.038t, and the coefficient of structure is 0.12.
(3) Compared with a single-stage rail-entering scheme based on the SABRE engine, the mass fraction of the effective load is improved from 3.74% to 5.89%, the emission cost of the effective load per unit mass can be effectively reduced, and the method is technically easy to implement.
Drawings
FIG. 1 shows thrust diagrams for the SABRE engine at different altitudes.
Fig. 2 shows a schematic diagram of the change of lift-drag ratio with angle of attack at different mach numbers.
FIG. 3 shows a force analysis diagram of a first-stage takeoff segment aircraft.
FIG. 4 shows a force analysis plot for a primary inspiratory mode segment.
Fig. 5 shows a force analysis diagram of the aircraft in the secondary climbing section.
Fig. 6 shows the acceleration as a function of time during the ascent of the aircraft.
Fig. 7 shows the speed variation over time during the ascent of the aircraft.
Fig. 8 shows the change in horizontal displacement over time during the ascent of the aircraft.
Fig. 9 shows the vertical displacement over time during the ascent of the aircraft.
Figure 10 shows a cross-sectional view of a two-stage on-track spacecraft in-track process.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings and specific preferred embodiments.
As shown in fig. 10, the method for the first-stage reusable two-stage orbit spacecraft comprises the following steps.
Step 1, constructing a model of the two-stage orbit spacecraft.
The total mass of the two-stage in-orbit spacecraft, also referred to herein as an aircraft, includes structural mass, propellant mass, and payload. And further obtaining the structural coefficient according to the structural quality of the aircraft.
The two-stage orbital spacecraft comprises a first stage and a second stage which are connected in series or in parallel, and the first stage and the second stage are preferably connected in parallel. The total takeoff mass of the aircraft in the present invention is preferably 345 t.
The first stage configuration is the trumpet creeper, which uses the SABRE engine as the power plant.
The first stage aerodynamic parameters are derived from a combination of CFD simulation and engineering methods, as shown in fig. 2. As can be seen from fig. 2, the maximum lift-drag ratio in the subsonic state is 9, the maximum lift-drag ratio in the mach number 2Ma is 5, which is lower than the maximum lift-drag ratio in the covey supersonic passenger aircraft 2Ma (the maximum lift-drag ratio in the covey supersonic passenger aircraft 2Ma is 7), which indicates that the aerodynamic parameters shown in fig. 2 are correct in the theoretical analysis level.
In addition, the structural coefficient of the first stage is preferably 0.1902, i.e., the structural mass of the first stage is 65.602 t.
The total consumption of the propellant can be obtained according to the working time of the engine in each stage and the flow rate of the propellant, and the total mass m of the propellant consumptiontThe calculation formula of (a) is as follows:
the total mass of propellant consumed in the first stage was 224.362t, as determined by equation (9).
The second stage is a small rocket, a liquid hydrogen/liquid oxygen rocket engine is used as a power device, and when the height of the rocket engine is higher than the design height of the spray pipe, the specific impulse is kept to be vacuum specific impulse and the numerical value is not changed any more; the vacuum specific impulse of the oxyhydrogen rocket is preferably 4500m/s, and the thrust of the rocket engine is controlled by adjusting the fuel flow during the flight.
The initial total mass of the second stage is preferably 55.038t, the structural coefficient of the second stage is preferably 0.12, i.e. the structural mass of the second stage is 6.605t, and the total mass of propellant consumed in the second stage is 28.128t, as calculated by equation (9). Thus, the payload mass in the second stage was found to be 20.305t, which is 5.89% of the total takeoff mass.
In the prior art, the take-off mass of a single-stage orbit aircraft of a Chinese trumpet creeper powered by an SABRE engine is 275t, the designed effective load is 10.275t, and the mass accounts for 3.74% of the total take-off mass. Therefore, compared with the prior art, the two-stage spacecraft launching method provided by the invention can obviously increase the mass (fraction) of the effective load, namely the mass fraction of the effective load is increased from 3.74% to 5.89%, meanwhile, the technical difficulty is greatly reduced, the method can be realized in a short time, and the method has great advantages.
Step 2, a first-stage takeoff section: the two-stage on-orbit spacecraft leaves the ground at a fixed elevation angle theta and at a speed of about Mach 0.5, and the flight time of the one-stage takeoff section is t1, and t1 is preferably 46 s. In the first-stage takeoff section, the aircraft is subjected to the action of gravity, aerodynamic force and ground friction resistance, and is accelerated horizontally under the action of the resultant external force, and the stress analysis is shown in fig. 3.
The equation of motion for the first-stage takeoff segment is expressed as follows:
in the formula: v represents the aircraft speed; t represents the thrust of the engine borne by the aircraft in the first-stage takeoff section; d represents the aerodynamic resistance of the aircraft; f represents the ground friction resistance borne by the aircraft; l represents the aerodynamic lift experienced by the aircraft; m represents the aircraft mass; m is0Representing an initial mass of the aircraft;representing an aircraft fuel consumption rate; ρ represents the atmospheric density of the aircraft at the position; s represents the aircraft characteristic area; cDRepresenting an aircraft drag coefficient; cLRepresenting the lift coefficient of the aircraft; f represents a ground friction coefficient; w represents the weight force borne by the aircraft; x represents the aircraft horizontal displacement.
The preferable values of the physical quantities are shown in table 1.
TABLE 1 first-level takeoff segment physical quantity measurement values
Step 3, a first-stage air suction mode section: after the primary takeoff section is finished, the aircraft climbs to the height of not less than 25km in an air suction mode under the action of gravity, aerodynamic force and thrust, the flight time of the primary air suction mode section is t2, and t2 is preferably 385 s.
Force analysis of the primary inspiratory mode segment as shown in fig. 4, the equation of motion of the primary inspiratory mode segment is expressed as follows:
in the formula: a isx、ayRespectively horizontal and verticalA directional acceleration; theta represents the aircraft angle of attack.
A is tox、ayExpressed as acos theta and asin theta, respectively, the above formula in formula (2) is multiplied by cos theta, and the following formula is multiplied by sin theta, so that simplification can be achieved
The above expression in the formula (2) is multiplied by sin theta, and the following expression is multiplied by cos theta, so as to simplify the process
By substituting formula (4) for formula (3)
In the formula:(g0r, H respectively representing the earth's surface gravitational acceleration, the earth's radius and the altitude at which the aircraft is located);
the differential equation of the motion of the climbing section of the air suction mode of the aircraft obtained by the formula (5) is
In the formula, T represents the thrust of an engine borne by the aircraft in a first-stage air suction mode section; m is1Representing the initial total mass of the aircraft in the primary air induction mode segment;representing an aircraft fuel consumption rate; g0Representing the earth's surface gravitational acceleration; k represents a lift-drag ratio coefficient; l represents the aerodynamic lift experienced by the aircraft; d represents the aerodynamic resistance of the aircraft; cLRepresenting the lift coefficient of the aircraft; cDRepresenting the aircraft drag coefficient.
In the formula (5), H (H.ltoreq.25 km) is a minute value relative to the earth radius R, and the approximation is consideredA simplified equation of motion for the first-order inspiratory mode segment is therefore:
in the first stage induction mode segment, as shown in FIG. 1, the thrust of the SABRE engine increases as altitude increases.
Step 4, a first-stage rocket mode section: after the first-stage air suction mode section is finished, the SABRE engine is switched to a rocket mode and pushes the two-stage in-orbit spacecraft to continuously climb, after the altitude reaches 100km, the first stage and the second stage are separated, and the first stage returns to the ground in an unpowered gliding mode to realize repeated use, such as a first-stage field returning section shown in figure 10.
The time of flight of the stage one rocket mode segment is t3, and t3 is preferably 304 s.
In the first stage rocket mode, the aircraft climbs at a fixed elevation angle, the stress analysis is the same as that in the first stage air suction mode, and the motion equation is shown in formula (6).
In the stage of the first rocket mode, the thrust of the SABRE engine remains unchanged, preferably 1458 kN.
Step 5, a secondary climbing section: after separation of the first and second stages, the second stage climbs vertically and makes a gravity turn at the end of the trajectory and feeds the payload into a 300km near-earth orbit at a speed in excess of 7700m/s, preferably 7925.009 m/s.
The flight time of the secondary climbing section is t4, t4 is preferably 33s, namely t2 > t3 > t1 > t4, and the total flight time is 768s
In the secondary climbing section, the atmospheric density is very thin due to the fact that the atmospheric density is above 100km, the aircraft climbs vertically, the effect of aerodynamic force on the aircraft is not considered any more, and the stress analysis is shown in fig. 5.
The equation of motion for the second-stage climb segment is expressed as follows:
in the formula: t represents the thrust of the engine borne by the aircraft in the secondary climbing section; w represents the weight force borne by the aircraft; m is2Representing the initial total mass of the aircraft when the rocket mode climbs to 100 km;represents a fuel consumption flow rate; isp represents the liquid oxygen/liquid hydrogen rocket engine specific impulse.
The mixing ratio of liquid oxygen/liquid hydrogen is 6.04 in the calculation process, and the required weight of the liquid oxygen or the liquid hydrogen can be quickly calculated according to the thrust of the engine borne by the aircraft in the secondary climbing section.
In step 5, the second stage climbs vertically and makes a gravity turn at the end of the trajectory and sends the payload into a 300km near-earth orbit at that speed.
The four-order Runge-Kutta method is adopted to solve the motion equations of the four stages, and the results are shown in FIGS. 6 to 9.
As can be seen in FIG. 6, during the first climb phase, the aircraft acceleration is low, with a maximum acceleration of 21.18m/s2, which is to reduce the work done to overcome the air resistance. In the second stage of climbing, the aircraft leaves the dense atmosphere, and the acceleration is rapidly increased so as to reach the rail entering speed.
As can be seen in FIG. 7, the first stage maximum velocity is 4882.076m/s and the second stage maximum velocity is 7909.964 m/s.
It can be seen from figure 8 that the first stage and second stage separation points, 1147.446km from the point of departure, are seen as more aerodynamic lift is utilized and the aircraft is flying a greater distance in the dense atmosphere.
As can be seen in FIG. 9, the first stage separates from the second stage at 100.01km, after which the second stage climbs rapidly, eventually reaching 300km height.
According to the simulation results, the total flight time is 768s, and the time used in each flight phase is shown in table 2.
TABLE 2 time periods corresponding to each flight phase
According to the simulation calculation result, when the aircraft reaches the 300km near-earth orbit, the speed is 7925.009m/s, the requirement of the orbit entering speed is met (the orbit entering speed is 7700m/s), and the two-stage orbit entering spacecraft orbit entering method is feasible.
Although the preferred embodiments of the present invention have been described in detail, the present invention is not limited to the details of the embodiments, and various equivalent modifications can be made within the technical spirit of the present invention, and the scope of the present invention is also within the scope of the present invention.
Claims (10)
1. A first-stage reusable two-stage orbit entering spacecraft orbit entering method is characterized in that: the method comprises the following steps:
step 1, constructing a model of a two-stage orbit spacecraft: the two-stage orbit spacecraft comprises a first stage and a second stage; the first-stage configuration is a trumpet creeper, and a SABRE engine is adopted as a power device; the second stage is a small rocket, and a liquid hydrogen/liquid oxygen rocket engine is used as a power device; the payload in the second stage accounts for 5.89% of the total takeoff mass;
step 2, a first-stage takeoff section: the two-stage in-orbit spacecraft leaves the ground at a fixed elevation angle theta from the vicinity of the equator of the earth, and the flight time of a first-stage takeoff section is t 1;
step 3, a first-stage air suction mode section: after the primary takeoff section is finished, under the action of gravity, aerodynamic force and thrust, climbing to the height of not less than 25km in an air suction mode, wherein the flight time of the primary air suction mode section is t 2;
step 4, a first-stage rocket mode section: after the first-stage air suction mode section is finished, the SABRE engine is switched to a rocket mode and pushes the two-stage in-orbit spacecraft to continuously climb, after the altitude reaches 100km, the first stage and the second stage are separated, and the first stage returns to the ground in an unpowered gliding mode to realize repeated use; the flight time of the first-stage rocket mode section is t 3;
step 5, a secondary climbing section: after the first and second stages are separated, the second stage climbs vertically and makes a gravity turn at the end of the trajectory and sends the payload into a 300km near-earth orbit at a speed exceeding 7700 m/s; the time of flight of the secondary climb segment is t 4.
2. The method of claim 1, wherein the first stage reusable two stage in-orbit spacecraft comprises: in step 1, the first level structural coefficient is 0.1902, and the second level structural coefficient is 0.12.
3. The method of claim 1, wherein the first stage reusable two stage in-orbit spacecraft comprises: in the step 3, when the SABRE engine is in the first-stage air suction mode section, the thrust of the SABRE engine is increased along with the rise of the height; and 4, keeping the thrust of the SABRE engine unchanged when the SABRE engine is in a first-stage rocket mode section.
4. The method of claim 1, wherein the first stage reusable two stage in-orbit spacecraft comprises: in step 2, the equation of motion of the first-stage takeoff section is expressed as follows:
in the formula: v represents the aircraft speed; t represents the thrust of the engine borne by the aircraft in the first-stage takeoff section; d represents the aerodynamic resistance of the aircraft; f represents the ground friction resistance borne by the aircraft; l represents the aerodynamic lift experienced by the aircraft; m represents the aircraft mass; m is0Representing an initial mass of the aircraft;representing an aircraft fuel consumption rate; ρ represents the atmospheric density of the aircraft at the position; s represents the aircraft characteristic area; cDRepresenting an aircraft drag coefficient; cLRepresenting the lift coefficient of the aircraft; f represents a ground friction coefficient; w represents the weight force borne by the aircraft; x represents the aircraft horizontal displacement.
5. The method of claim 1, wherein the first stage reusable two stage in-orbit spacecraft comprises: in step 3, the equation of motion of the first-stage inspiration mode segment is expressed as follows:
wherein,
in the formula, T represents the thrust of an engine borne by the aircraft in a first-stage air suction mode section; m is1Representing the initial total mass of the aircraft in the first-stage suction mode segmentRepresenting an aircraft fuel consumption rate; g0Representing the earth's surface gravitational acceleration; k represents a lift-drag ratio coefficient; l represents the aerodynamic lift experienced by the aircraft; d represents the aerodynamic resistance of the aircraft; cLRepresenting the lift coefficient of the aircraft; cDRepresenting the aircraft drag coefficient.
6. The method of claim 1, wherein the first stage reusable two stage in-orbit spacecraft comprises: in step 4, the equation of motion of the first-stage rocket mode section is expressed as follows:
wherein,
in the formula, T represents the engine thrust borne by the aircraft in the first-stage rocket mode section; m is1Representing the initial total mass of the aircraft in a stage one rocket mode segment;representing an aircraft fuel consumption rate; g0R, H respectively representing the acceleration of gravity on the earth's surface, the radius of the earth, and the altitude at which the aircraft is located; k represents a lift-drag ratio coefficient; l represents the aerodynamic lift of the aircraft; d represents the aerodynamic resistance of the aircraft; cLRepresenting the lift coefficient of the aircraft; cDRepresenting an aircraft drag coefficient; theta represents the aircraft angle of attack.
7. The method of claim 1, wherein the first stage reusable two stage in-orbit spacecraft comprises: in step 5, the equation of motion of the secondary climbing section is expressed as follows:
in the formula: t represents the thrust of the engine borne by the aircraft in the secondary climbing section; w represents the weight force borne by the aircraft; m is2Representing the initial total mass of the aircraft when the rocket mode climbs to 100 km;represents a fuel consumption flow rate; isp represents the liquid oxygen/liquid hydrogen rocket engine specific impulse.
8. The method of claim 1, wherein the first stage reusable two stage in-orbit spacecraft comprises: t 1-46 s, t 2-385 s, t 3-304 s, t 4-33 s, total time of flight 768 s.
9. The method of claim 1, wherein the first stage reusable two stage in-orbit spacecraft comprises: in step 5, the second stage climbs vertically and makes a gravity turn at the end of the trajectory and sends the payload into a 300km near-earth orbit at 7925.009 m/s.
10. The method of claim 1, wherein the first stage reusable two stage in-orbit spacecraft comprises: in the step 1, a liquid hydrogen/liquid oxygen rocket engine is adopted as a power device in the second stage, and when the height of the rocket engine is higher than the design height of a spray pipe, the specific impulse is kept to be vacuum specific impulse and the numerical value is not changed any more; the vacuum specific impulse of the oxyhydrogen rocket is 4500m/s, and the thrust of the rocket engine is controlled by adjusting the fuel flow in the flight process.
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Cited By (3)
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|---|---|---|---|---|
| CN109398762A (en) * | 2018-10-17 | 2019-03-01 | 湖北航天技术研究院总体设计所 | A kind of solid-rocket enters rail ballistic design method based on elliptical transfer orbit |
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