GB2388578A - Satellite launch vehicles and a method of launching satellites into orbit - Google Patents
Satellite launch vehicles and a method of launching satellites into orbit Download PDFInfo
- Publication number
- GB2388578A GB2388578A GB0211505A GB0211505A GB2388578A GB 2388578 A GB2388578 A GB 2388578A GB 0211505 A GB0211505 A GB 0211505A GB 0211505 A GB0211505 A GB 0211505A GB 2388578 A GB2388578 A GB 2388578A
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- Prior art keywords
- stage
- orbit
- arcing
- profile
- launch vehicle
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- 238000000034 method Methods 0.000 title claims description 10
- 230000001174 ascending effect Effects 0.000 claims abstract description 5
- 238000012546 transfer Methods 0.000 description 9
- 238000002347 injection Methods 0.000 description 6
- 239000007924 injection Substances 0.000 description 6
- 230000009194 climbing Effects 0.000 description 3
- 230000000694 effects Effects 0.000 description 3
- 239000000446 fuel Substances 0.000 description 3
- 238000013461 design Methods 0.000 description 2
- 239000007788 liquid Substances 0.000 description 2
- 239000004449 solid propellant Substances 0.000 description 2
- 238000004458 analytical method Methods 0.000 description 1
- 238000013459 approach Methods 0.000 description 1
- 230000005540 biological transmission Effects 0.000 description 1
- 238000007796 conventional method Methods 0.000 description 1
- 230000005484 gravity Effects 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 239000003380 propellant Substances 0.000 description 1
- 230000009291 secondary effect Effects 0.000 description 1
- 239000000243 solution Substances 0.000 description 1
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/002—Launch systems
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/242—Orbits and trajectories
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/242—Orbits and trajectories
- B64G1/2427—Transfer orbits
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- Engineering & Computer Science (AREA)
- Remote Sensing (AREA)
- Aviation & Aerospace Engineering (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Radar, Positioning & Navigation (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
A multi-stage launch vehicle for delivering a satellite into orbit, configured to follow a trajectory which comprises at least one lower stage, and an upper stage arranged to apply thrust over an arcing flight profile comprising an initially ascending, then descending arc. The descending arc preferably comprises at least a quarter of the arcing profile. The arcing profile itself preferably experiences a net descent. Further, there are preferably three lower stages with the upper stage preferably arranged to re-apply thrust after a coasting period.
Description
, 2388578
- 1 SATELLITE LAUNCH VEHICLES AND A METIIOD OF LAUNCHING
SATELLITES
This invention relates to satellite launch vehicles, and to a method of launching satellites.
5 The trajectory resulting from a typical satellite launch is shown in Figure 1. This launch vehicle has four stages Stages 1-3 take the launch vehicle to position A (an altitude of typically 100 to 200km), and Stage 4 takes it to position B (an altitude of typically 200 to 300km). At position B. the launch vehicle has picked up sufficient velocity to orbit the Earth in a transfer elliptical to orbit C. The centre of the Earth lies at one of the foci of the ellipse. The elliptical orbit it used to transfer the spacecraft to a bigger orbit, in which it remains by virtue of an injection burn at position D (an altitude of several hundred kilometres). Both the burn from positions A to B and at position D are provided by the fourth stage of the launch vehicle.
s The altitude of the launch vehicle from the Earth increases continuously up to position A, and beyond position A up to position B. Figure 2 shows the altitude profile over the fourth stage. Beyond position B. the spacecraft coasts until it receives its injection burn at position D. The Applicants have considered ways of improving the overall efficiency 20 of this known flight profile.
The invention provides a multi-stage launch vehicle for delivering a satellite into orbit, which comprises at least one lower stage, and an upper stage arranged to apply thrust over an arcing flight profile comprising an initially ascending, then descending arc.
:s The invention also provides a method of utilising a multi-stage launch vehicle to deliver a satellite into orbit, which comprises the steps of providing thrust from at least one lower stage, and providing thrust from an upper stage over an arcing flight profile comprising an initially ascending, then descending arc.
( It has been found that the energy input to the spacecraft can be increased if part of the thrust takes place over an arcing flight profile, compared to the previous strategy of thrusting with continuously increasing height.
The arcing profile may advantageously be arranged so that the launch 5 vehicle experiences a net descent over the arcing profile.
There may be three lower stages. The upper stage may itself comprise more than one stage.
The invention is particularly applicable to upper stages having a ratio of thrust to weight of significantly less than 19.
to The invention will now be described, by way of example, with reference to the accompanying drawings, in which: Figure 1 illustrates a known launch strategy; Figure 2 shows an altitude history for the known launch strategy of Figure 1; s Figure 3 illustrates a launch strategy in accordance with the invention with the thrusVweight (TfN) ratio of an upper stage of the launch vehicle approximately equal to 0.2; Figure 4 shows an altitude history for the launch strategy of Figure 3, showing the corresponding applications of thrust; and no Figure 5 shows an altitude history for a launch strategy in accordance with the invention with the thrusVweight (TNV) ratio of an upper stage of the launch vehicle approximately equal to 0.4.
The dimensions in Figures 1 and 3 are not to scale. In fact, the altitude has been multiplied by five with respect to the radius of the Earth in order to 25 illustrate trajectory shape more clearly.
Referring to Figures 1 and 2, the launch vehicle has three lower stages and one upper stage. Traditional launcher ascent trajectories tend to follow a similar profile. A vertical take-off is followed by gradual pitch-over as the launcher picks up speed. The three lower stages take the launch vehicle above
( - 3 the approximate level of the atmosphere (the broken line at 100km approx.) and up to position A (between 100-200km).
The fourth stage burns over two separate periods of time. Firstly, it burns immediately after position A has been reached and until position B has 5 been reached, in order to provide the spacecraft with sufficient velocity to attain elliptical orbit C, of which the centre of the Earth lies at one of the foci. The second part of the fourth stage bum is to circularise the orbit at the apogee of the elliptical transfer orbit at position D, or alternatively to place the vehicle in a new elliptical orbit.
to The flight profile is such that the spacecraft is climbing and only at injection at position D does the profile level off. Of course, the vehicle is not powered between positions B and D and is coasting. In some cases, a limited degree of descent is experienced near the end of the upper stage burn at position B. prior to injection at the transfer orbit perigee.
Typical launch vehicles have a low thrusVweight (TNV) ratio for the upper stage. In particular, small launch vehicles tend to utilise even lower thrusVweight (TAN) levels. The problem with such an upper stage, following a traditional flight profile, is the lack of efficiency in transferring energy to the orbit.
This arises from the extended duration of the burn, which means that energy is no not transferred at the optimum location. The first of the upper stage burns is much longer than the second and experiences the greatest energy loss.
Referring to Figures 3 and 4, a multi-stage launch vehicle in accordance with the invention has three lower stages. A vertical take-off is followed by a gradual pitch-over as the launcher picks up speed. The third stage cuts out at 25 position A in Figures 3 and 4 (an altitude of typically between 100 to 200km).
The fourth stage has two burns. The first from position A to position B is to generate sufficient velocity to orbit in transfer orbit C. The second burn of the upper stage is at position D to circularise the elliptical orbit, or to place the craft in a different elliptical orbit. Without this injection burn, the spacecraft would 30 continue on an elliptical path, with the Earth at one of the foci of the ellipse, and pass through position B again.
- 4 The basis of the strategy is to force the upper stage to thrust over an initially climbing, then descending arc. The trajectory exhibits a significant arcing profile. In Figures 3 and 4, the spacecraft descends, while thrusting, for approximately 200km, after the upper stage reaches its maximum altitude during the first burn. Such an arc is more of a Ballistic' profile than traditional trajectories. The result of this is that the perigee of the elliptical orbit C is nearer to the Earth than the traditional launch trajectory of Figure 1 (at approximately 120km).
This is still sufficiently high to avoid problems with atmospheric encounter.
to Energy is transferred more efficiently to the spacecraft than hitherto for the following reason. The energy converted by the rocket motor into energy of the spacecraft is equal to the work done in traversing from position A to position B. This is the product of the force (i.e. the thrust) and the distance travelied.
The distance travailed is the integral of the velocity of the spacecraft with 15 respect to time. The spacecraft picks up speed from position A to position B. as it does with the traditional launch trajectory of Figure 1, but there is also a net interchange of potential (gravitational) and kinetic energy compared with the strategy of Figure 2, because the altitude of position B is less than that of position A. The speed at B is therefore increased, and the average velocity 20 between positions A and B is also increased compared to the traditional strategy. Given that the thrust is applied over a higher average velocity, more energy is transferred to the spacecraft.
There is a loss of energy at the injection burn at position D in Figure 2 compared to Figure 1, but this is more than offset by the energy gain referred 25 to. It is not even essential for the upper stage to ignite just as the third stage cuts off. If desired, the launch vehicle could be allowed to coast after the third stage is cut off until it begins to descend, and the fourth stage could be ignited at that point. In this case, the perigee of the transfer orbit would be nearer to 30 the Earth and losses due to the atmosphere would reduce the energy gain associated with the fourth stage.
( - 5 The launch vehicle and the method of using it will now be described in more detail. The optimum strategy can be described by the following sequence.
In the third (or penultimate) stage cut-off, the launch vehicle is directed into a shallow climb (the exact value, with corresponding altitude/speed being 5 optimised for each case). The fourth (or final) stage is ignited immediately after the preceding stage has been cut off. A low incidence value is maintained (the exact profile being optimised in each case). The incidence is the angle between the thrust axis and the velocity vector of the launch vehicle. A low incidence value allows the onset of a descending phase after an initial climb. The upper to engine stage is cut off during the descending phase. The exact altitude and flight path at engine cut-off is determined on a case-by-case basis. Engine cut off could be as late as perigee (i.e. where the descending phase terminates), or in certain cases after perigee on the subsequent climbing phase of the trajectory. The optimum engine cut off altitude is related to the speed required at engine cut off (i.e. the energy of the parking orbit reflected in the parking orbit apogee altitude) and also the speed at upper stage ignition Net altitude reduction may be experienced during the upper stage burn, and this is illustrated in Figure 5.
Regarding the coast phase, this phase passes through the perigee if the no upper stage cut off is pre-perigee. The perigee may be as low as 100km. The value depends on heating rate constraints that the upper stage of the spacecraft can tolerate and also on operational constraints. It may be necessary to constrain the perigee altitude to a minimum allowed value exceeding 1 OOkm (for example 120km). After the perigee the stage climbs to a target parking orbit s apogee altitude. This will be at an altitude close to the required circular orbit altitude or elliptical orbit apogee or perigee altitude. Shortly before apogee is reached, the engines are reignited to increase speed, leaving the vehicle in the desired circular or elliptical orbit.
This strategy is particularly applicable where the thrusVweight (TNV) ratio 30 of the upper stage is low. This is typically the case for small launch vehicles, where TNY often takes a value of less than 19. The gains obtained by the optimised strategy become increasingly significant in comparison with the more
- 6 conventional method as lower TN is considered. Conversely, a stage with TNV in the region of 19 is unlikely to experience significant benefit.
The secondary effect of increased efficiency with lower TNV is that it offers the possibility to change the design of the upper stage. A stage utilising the optimised strategy may be designed with lower thrust engines, thus making a saving on engine/structure mass which may then be used for payload.
Therefore the benefit from this type of strategy can be amplified beyond the simple fuel saving on a case by case basis.
It is possible to analyse the means by which the strategy gains in terms to of fuel efficiency, in comparison with the more traditional upper stage profile.
For a small launch vehicle, fuel savings of typically 2% of the fully fuelled upper stage mass (depending on the thrusVweight (TNV) ratio) can be obtained as a direct result of the optimised upper stage strategy. This is significant when compared with a total launcher payload mass in the region of 25% of the total s upper stage mass. The ability to efficiently inject upper stages with low thrusVweight (TNV) ratios into orbit allows the evolution of their design to exploit the mass savings from using lower thrust components.
The fundamental equation governing the transmission of energy from the propulsion system to the spacecraft is as follows: 20 AK= JForce.Vd' (11) ' Where is the useful work done to the spacecrafVfuel system by the propulsion system thrusting for a period T. V is the spacecraft velocity vector, and Force is the applied thrust vector.
25 The equation illustrates two important points. The energy transfer is maximised if the magnitude of the velocity vector is at a high value. The speed within any orbit takes its maximum value at the point of closest approach to the Earth. Using a parking orbit with a low perigee and ensuring that the launch vehicle injects into the parking orbit at an altitude close to perigee is a means to
- 7 ensure that speed is at a high value. The energy transfer is maximised if 'thrust vector losses' (i.e. the cosine effect due to the angle between the thrust axis direction and the velocity vector direction) are minimised. In other words, the angle should be maintained at a value as close to zero as possible.
The gains of the optimised strategy thus presented should be considered in the following context. In arriving at an optimum profile, a series of tradeoffs take place. By following an arcing trajectory, it is possible to maintain low values of incidence throughout the upper stage bum. A zero incidence trajectory will follow an arc with the degree of fuming influenced entirely by o gravity.
By maintaining a low value of incidence angle, the 'cosine loss' is minimised. The extent of climb or descent experienced during the upper stage engine burn arc is significantly effected by: (1) the value of the flight path angle at the termination of the preceding stage and (2) the extent of any "thrust vectoring" i.e. incidence greater than zero, steering the trajectory. This can either increase or decrease the altitude range of the arc in comparison with the zero incidence case.
zo if the arc is too high, the efficiency of energy transfer is reduced due to the reduced speed. To maintain a trajectory with a steadily increasing altitude profile requires greater incidence to offset the natural gravitational effect of the trajectory. A complication is the effect of the atmosphere in the region of the 25 subsequent perigee. The optimised value in the cases examined was in the region of 100km in altitude. At this altitude, the spacecraft experiences a small atmospheric drag. This tends to absorb the spacecraft energy. Therefore there is a trade-off between the desire to maximise the extent of the descending phase of the upper stage bum and the need to maintain a perigee altitude that so is sufficiently high to prevent significant energy loss to the atmosphere.
- 8 The optimal solution to this tradeoff is the significant arcing trajectories described herein.
The lower stages may use liquid or solid propellant. The upper stage may use liquid propellant or solid propellant. The upper stage may itself be in two stages from, for example, A to B. Variations may of course be made without departing from the scope of the invention. Thus, there may be more or less than three lower stages.
Claims (12)
1. A multi-stage launch vehicle for delivering a satellite into orbit, which comprises at least one lower stage, and an upper stage arranged to apply thrust over an arcing flight profile comprising an initially ascending, then descending arc.
2. A launch vehicle as claimed in Claim 1, in which the descending arc
comprises at least a quarter of the arcing profile.
3. A launch vehicle as claimed in Claim 1 or Claim 2, in which the arcing flight profile experiences a net descent.
to
4. A multi-stage launch vehicle as claimed in any one of Claims 1 to 3, in which there are three lower stages.
5. A multi-stage launch vehicle as claimed in any one of Claims 1 to 4, in which the upper stage is also arranged to re-apply thrust after a coasting period, in order to inject the payload into a desired circular or elliptical 5 orbit.
6. A multi-stage launch vehicle substantially as herein described with reference to Figures 3 to 5 of the accompanying drawings.
7. A method of utilising a multi-stage launch vehicle to deliver a satellite into orbit, which comprises the steps of providing thrust from at least one no lower stage, and providing thrust from an upper stage over an arcing flight profile comprising an initially ascending, then descending arc.
8. A method as claimed in Claim 7, in which the descending arc comprises at least a quarter of the arcing profile.
9. A method as claimed in Claim 7 or Maim 8, in which the arcing flight 25 profile experiences a net descent.
10. A method as claimed in any of Claims 7 to 9, in which there are three lower stages.
- 10
11. -A method as claimed in any of Claims 7 to 10, in which the upper stage re-applies thrust after a coasting period, to inject the payload into a desired circular or elliptical orbit.
12. A method substantially as herein described with reference to Figures 3 to 5 5 of the accompanying drawings.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB0211505A GB2388578A (en) | 2002-05-18 | 2002-05-18 | Satellite launch vehicles and a method of launching satellites into orbit |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB0211505A GB2388578A (en) | 2002-05-18 | 2002-05-18 | Satellite launch vehicles and a method of launching satellites into orbit |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| GB0211505D0 GB0211505D0 (en) | 2002-06-26 |
| GB2388578A true GB2388578A (en) | 2003-11-19 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB0211505A Withdrawn GB2388578A (en) | 2002-05-18 | 2002-05-18 | Satellite launch vehicles and a method of launching satellites into orbit |
Country Status (1)
| Country | Link |
|---|---|
| GB (1) | GB2388578A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US11377237B1 (en) * | 2019-05-01 | 2022-07-05 | United Launch Alliance, L.L.C. | Orbital rendezvous techniques |
Families Citing this family (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110775301A (en) * | 2019-12-04 | 2020-02-11 | 哈尔滨工业大学 | Aircraft with high rail-entering efficiency and strong maneuvering capability and rail-entering method thereof |
| CN112393835B (en) * | 2020-11-03 | 2021-12-21 | 西北工业大学深圳研究院 | Small satellite on-orbit thrust calibration method based on extended Kalman filtering |
Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0640524A1 (en) * | 1993-08-24 | 1995-03-01 | Orbital Sciences Corporation | Method for injecting payloads into orbit |
-
2002
- 2002-05-18 GB GB0211505A patent/GB2388578A/en not_active Withdrawn
Patent Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0640524A1 (en) * | 1993-08-24 | 1995-03-01 | Orbital Sciences Corporation | Method for injecting payloads into orbit |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US11377237B1 (en) * | 2019-05-01 | 2022-07-05 | United Launch Alliance, L.L.C. | Orbital rendezvous techniques |
Also Published As
| Publication number | Publication date |
|---|---|
| GB0211505D0 (en) | 2002-06-26 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| WAP | Application withdrawn, taken to be withdrawn or refused ** after publication under section 16(1) |