CN107013334B - A kind of double combustion chamber's Scramjet Inlet and air intake control method - Google Patents
A kind of double combustion chamber's Scramjet Inlet and air intake control method Download PDFInfo
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- CN107013334B CN107013334B CN201710086436.0A CN201710086436A CN107013334B CN 107013334 B CN107013334 B CN 107013334B CN 201710086436 A CN201710086436 A CN 201710086436A CN 107013334 B CN107013334 B CN 107013334B
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
- F02C7/057—Control or regulation
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
- F02C7/042—Air intakes for gas-turbine plants or jet-propulsion plants having variable geometry
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
- F02K7/10—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
- F02K7/14—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines with external combustion, e.g. scram-jet engines
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- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
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Abstract
It is an object of the invention to overcome the shortage of prior art, a kind of double combustion chamber's Scramjet Inlet and air intake control method are provided, solves the technical barrier of wide scope non-axis symmetry double combustion chamber's scramjet engine air inlet.The air intake duct uses two dimensional inlet configuration, and sub- combustion runner and super burn runner are divided by support plate along flow direction, the sub- combustion channel is at least one, the super burn runner at least two and divide equally in Asia combustion runner both sides, by increasing contract section adjustment type face and expansion segment adjustment type face in each runner design, to control inlet throat size, realize needs of each combustion chamber to different compression degree air, it widens engine work range of Mach numbers, improve engine performance, solve the problems, such as difficult with turbogenerator relay.
Description
Technical field
The invention belongs to aerospace technical field of engines, are related to a kind of double combustion chamber's Scramjet Inlet
And air intake control method.
Background technology
With the fast development of hypersonic technology, near space has become the New Frontier that great powers in the world competitively fight for,
As the new strategic high ground of national security, near space hypersonic aircraft is increasingly becoming each technologically advanced national national defence and builds
If pay close attention to direction.Near space hypersonic aircraft is using air suction type combined engine as power, horizontal can be risen
Drop, in dense atmosphere, near space and the reusable hypersonic aircraft of LEO maneuvering flight.
In a variety of air suction type combined engines, the assembly power scheme of turbogenerator and wide scope punching engine is most
Future in engineering applications, but " thrust wide gap " is the biggest problem existing for the program.Current wide scope punching engine mainly uses
Dual-mode Scramjet, work Mach number lower limit is Ma4 or so, and existing shelf turbogenerator work Mach number
In Ma2.5 or so, the two is difficult to realize smoothly be connected the upper limit.The main means for solving " thrust wide gap " at present have：First, it uses
High-speed turbine technology and the water spray precooling skill upgrading turbogenerator work Mach number upper limit, but high-speed turbine technical difficulty pole
Height, the U.S. is by the still difficult breakthrough that studies for a long period of time, and pre- refrigeration technique of spraying water need to load a large amount of water and increase structure complexity；Second is that
It introduces rocket engine and makes up thrust deficiency, the program so that combined engine system is more complicated, and needs a variety of propellant combustions
Material.
In a variety of air suction type combined engines, the assembly power scheme of turbogenerator and wide scope punching engine is most
Future in engineering applications, but " thrust wide gap " is the biggest problem existing for the program.Current wide scope punching engine mainly uses
Dual-mode Scramjet, work Mach number lower limit is Ma4 or so, and existing shelf turbogenerator work Mach number
In Ma2.5 or so, the two is difficult to realize smoothly be connected the upper limit.The main means for solving " thrust wide gap " at present have：First, it uses
High-speed turbine technology and the water spray precooling skill upgrading turbogenerator work Mach number upper limit, but high-speed turbine technical difficulty pole
Height, the U.S. is by the still difficult breakthrough that studies for a long period of time, and pre- refrigeration technique of spraying water need to load a large amount of water and increase structure complexity；Second is that
It introduces rocket engine and makes up thrust deficiency, the program so that combined engine system is more complicated, and needs a variety of propellant combustions
Material.
Double combustion chamber's scramjet engine is lower than Dual-mode Scramjet work Mach number lower limit, is
Ma3.3 or so, by adjustable air intake technology, work Mach number lower limit can be opened up further down to Ma2.5 or so, therefore turbine is sent out
The assembled scheme of motivation and double combustion chamber's scramjet engine is more conducive to solve " thrust wide gap " problem, based on shelf turbine
The assembly power scheme of engine is more feasible.
Existing double combustion chamber's scramjet engine is axial symmetry configuration, using axialsymmetrical inlet, is more suitable for axis pair
Claim layout, be difficult to realize be arranged in parallel with turbine channel in terms of runner and structure design and share gas handling system, and existing
Technology is using fixed geometry air intake duct, it is difficult to adapt to wider work range of Mach numbers.
Invention content
It is an object of the invention to overcome the shortage of prior art, a kind of double combustion chamber's scramjet engine air inlet is provided
Road solves the technical barrier of wide scope non-axis symmetry double combustion chamber's scramjet engine air inlet.
The technical solution of the present invention：
A kind of double combustion chamber's Scramjet Inlet, the air intake duct use two dimensional inlet configuration, and along stream
Sub- combustion runner and super burn runner are divided by support plate, the sub- combustion channel is at least one, and the super burn runner is at least two
It is a and divide equally in Asia combustion runner both sides；
The sub- combustion is consistent with super burn flow passage structure, is made of fixed face, adjustable type face and connects hinge, wherein, institute
State fixed face include external pressure miniature face, contract segment type face, venturi type face, expansion segment type face, lip type face and with burning
Room connecting-type face；The adjustable type face includes contract section adjustment type face and expansion segment adjustment type face, and the contract section is adjusted
The one end in type face and expansion segment adjustment type face is individually fixed in contract section starting point and expansion segment endpoint, and the other end is respectively activity
End, the connects hinge include hinge a and b, and the hinge a and b is located at contract section starting point and expansion segment endpoint, and institute respectively
Contract section adjustment type face and expansion segment adjustment type face is stated to be revolved by design rotation angle respectively around the hinge a and b
Turn.
Further, during rotation is adjusted, the activity in the contract section adjustment type face and expansion segment adjustment type face
The vertical range apart from lip type face is held to remain equal.
Further, after the contract section adjustment type face and expansion segment adjustment type face rotate, the contract section is adjusted
Type face, contract segment type face, venturi type face, expansion segment type face and expansion segment adjustment type face surround a cavity, the shape in cavity
Into pneumatic venturi type face of the vortex as air intake duct runner, the corresponding venturi in the pneumatic venturi type face is highly defined as adjustable type face
Runner venturi height H after rotationth, obtained by formula (1)：
Wherein, HthRunner venturi height after being rotated for adjustable type face, HcHeight is captured for air intake duct runner, q (Ma) is stream
Flow function, Ma0And MathRespectively free stream Mach number and runner venturi Mach number,For discharge coefficient, σthFor runner venturi stagnation pressure
Recovery coefficient.
Further, the rotation angle includes the rotation angle θ in contract section adjustment type face7With expansion segment adjustment type face
Rotation angle θ8, obtained by formula (2)；
Wherein, H7For contract section adjustment type face fixing end and lip type face vertical height, H8Consolidate for expansion segment adjustment type face
Fixed end and lip type face vertical height, HthRunner venturi height after being rotated for adjustable type face, L7For contract section adjustment type face two
Distance between endpoint, L8The distance between the two-end-point of expansion segment adjustment type face, θ2For contract segment type face two end point connecting line and venturi type face
Between angle (taking acute angle), θ4For the angle (taking acute angle) between expansion segment type face two end point connecting line and venturi type face.
The rotation angle θ7And θ8When being 0 °, the contract section adjustment type face and expansion segment adjustment type face respectively with it is interior
It shrinks segment type face and expansion segment type face paste is closed；
The rotation angle θ7And θ8During for maximum value, the contract section adjustment type face and expansion segment adjustment type face are complete
Close the runner.
Further, it is as follows to fix quasi spline for the air intake duct：It sets air intake duct and works range of Mach numbers as Mamin~
Mamax, the air intake duct fixed face is according to MamaxThe requirement of the performance indicators such as discharge coefficient, total pressure recovery and MaminRise
Dynamic performance requirement is designed and is obtained, this technology being known in the art.
Further, when free stream Mach number is more than MaminWhen, according to engine performance demand and the tune of Design of Inlet
Rule is saved, the hinge control adjustable type face rotation is adjusted, until free stream Mach number is Mamax。
The present invention also provides a kind of above-mentioned air intake control method of double combustion chamber's Scramjet Inlet, including with
Lower process：
When flight Mach number reaches the first Mach number, the contract section adjustment type face of the sub- combustion runner and super burn runner
It is close to contract segment type face and expansion segment type face respectively with expansion segment adjustment type face, subsonic combustion room and supersonic speed combustion chamber are opened
Initial point fire, the subsonic combustion room and supersonic speed combustion chamber are sub- combustion mode；
When flight Mach number reaches the second Mach number, the contract section adjustment type face of the sub- combustion runner and super burn runner and
Expansion segment adjustment type face proceeds by rotation, with the increase of flight Mach number, controls the interior receipts of sub- combustion runner and super burn runner
Contracting section adjustment type face and the rotation of expansion segment adjustment type face so that the subsonic combustion room and supersonic speed combustion chamber are in sub- combustion
Mode；
When flight Mach number reaches third Mach number, the contract section adjustment type face of super burn runner and expansion segment adjustment type face
It stops rotating adjusting, with the increase of flight Mach number, supersonic speed combustion chamber gradually changes from Asia combustion mode to super burn mode；
When flight Mach number reaches the 4th Mach number, supersonic speed combustion chamber is in super burn mode, completes mode transformation, with
The increase of flight Mach number, the contract section adjustment type face and expansion segment adjustment type face of super burn runner still stop, and do not rotate adjusting；
When flight Mach number reaches the 5th Mach number, the contract section adjustment type face of super burn runner and expansion segment adjustment type face
Start rotation to adjust, with the increase of flight Mach number so that the supersonic speed combustion chamber is in super burn mode always；
When flight Mach number reaches the 6th Mach number, the contract section adjustment type face of the sub- combustion runner and super burn runner and
Expansion segment adjustment type face is stopped rotating adjusting, and always in sub- combustion mode, supersonic speed combustion chamber begins for shown subsonic combustion room
It is in super burn mode eventually；
Between second Mach number and the 6th Mach number, contract section adjustment type face and the expansion segment of sub- combustion runner are controlled
Adjustment type face rotates so that the subsonic combustion room is always in sub- combustion mode.
Beneficial effects of the present invention：
1st, by the way of two dimensional inlet runner, double combustion chamber's scramjet engine non-axis symmetry configuration is solved
Air inlet problem, and can preferably be applied to parallel turbine base combined engine scheme.
2nd, it adjusts using geometry and pneumatically adjusts the regulation measure being combined and preferably resolve double combustion chamber's ultra-combustion ramjet
Engine wide scope work problem helps to widen engine work Mach number lower limit, while there is the regulation scheme mitigation to adjust
It saves additional mass, reduce many advantages, such as high temperature dynamic sealing demand.
3rd, air intake duct subsonic combustion room runner and supersonic speed combustion chamber runner are separately adjustable, and air intake duct is sub- to entering
The compression degree of the air-flow of velocity of sound combustion chamber is more than the air-flow into supersonic speed combustion chamber, is conducive to match the respective need of combustion chamber
It asks, convenient for burning tissues, improves engine performance.
4th, it realizes intake valve deposit using the existing air inlet adjustment function of combined engine, widens double combustion chamber's ultra-combustion ramjet
The work range of Mach numbers of engine, without additionally increasing regulating mechanism.
Description of the drawings
Included attached drawing is used for providing being further understood from the embodiment of the present invention, and which constitute one of specification
Point, for illustrating the embodiment of the present invention, and come together with word description to illustrate the principle of the present invention.It should be evident that below
Attached drawing in description is only some embodiments of the present invention, for those of ordinary skill in the art, is not paying creation
Property labour under the premise of, can also be obtained according to these attached drawings other attached drawings.
Fig. 1 is the aloof ramjet engine air inlet runner schematic layout pattern in double combustion chamber provided in an embodiment of the present invention；
Fig. 2 is the aloof punching engine flow path adjustment scheme schematic diagram in double combustion chamber in the embodiment of the present invention；
Fig. 3 is the aloof punching engine flow path adjustment parameter declaration schematic diagram in double combustion chamber in the embodiment of the present invention.
In figure, A, super burn runner；B, sub- combustion runner；C, super burn runner；1st, external pressure miniature face；2nd, contract segment type face；3、
Venturi type face；4th, expansion segment type face；5th, lip type face；6th, with combustion chamber connecting-type face；7th, contract section adjustment type face；8th, it expands
Section adjustment type face；9th, hinge a；10th, hinge b；H7For 7 fixing end of contract section adjustment type face and 5 vertical height of lip type face；H8
For 8 fixing end of expansion segment adjustment type face and 5 vertical height of lip type face；HthFor the required venturi height adjusted；θ2For interior receipts
Angle (taking acute angle) between 2 two end point connecting line of contracting segment type face and venturi type face 3；θ4For 4 two end point connecting line of expansion segment type face with
Angle (taking acute angle) between venturi type face 3.
Specific embodiment
Specific embodiments of the present invention are described in detail below in conjunction with attached drawing.In the following description, for solution
Purpose and not restrictive is released, elaborates detail, to help to be apparent from the present invention.However, to people in the art
It will be apparent that the present invention can also be put into practice in the other embodiments departing from these details for member.
It should be noted that in order to avoid because having obscured the present invention during unnecessary details, only show in the accompanying drawings
The device structure closely related with scheme according to the present invention and/or processing step are gone out, and have been omitted with relationship of the present invention not
Big other details.
Embodiment 1
The present embodiment provides a kind of aloof ramjet engine air inlet in double combustion chamber, Fig. 1 is provided in an embodiment of the present invention
The aloof ramjet engine air inlet runner schematic layout pattern in double combustion chamber；Fig. 2 is air intake duct flow path adjustment in the embodiment of the present invention
Scheme schematic diagram；Fig. 3 illustrates flow path adjustment parameter declaration schematic diagram.The air intake duct uses two dimensional inlet configuration, and edge
Flow direction is divided into sub- combustion runner B and super burn runner A and C by support plate, and the B runners are intermediate flow channel, connect subsonic combustion room,
A, C runner is distributed in B runners both sides connection supersonic speed combustion chamber.
Further, it is arranged by engine blockization, above-mentioned runner layout can also be that AABAA, ABABC, ABBC etc. are more
Kind combining form.
The arbitrary runner design of A, B, C is as follows：
The runner is made of fixed face, adjustable type face and connects hinge, wherein, the fixed face includes external compression
Type face 1, contract segment type face 2, venturi type face 3, expansion segment type face 4, lip type face 5 and with combustion chamber connecting-type face 6；It is described
Adjustable type face includes contract section adjustment type face 7 and expansion segment adjustment type face 8, the contract section adjustment type face 7 and expansion segment
The one end in adjustment type face 8 is individually fixed in contract section starting point and expansion segment endpoint, and the other end is respectively movable end, the connection
Hinge includes hinge a9 and b10, and the hinge a9 and b10 is located at contract section starting point and expansion segment segment endpoint, and described respectively
Contract section adjustment type face 7 and expansion segment adjustment type face 8 can be carried out respectively around the hinge a9 and b10 by design rotation angle
Rotation.
Further, it is as follows to fix quasi spline for the air intake duct：It sets air intake duct and works range of Mach numbers as Mamin~
Mamax, the air intake duct fixed face is according to MamaxThe requirement of the performance indicators such as discharge coefficient, total pressure recovery and MaminRise
Dynamic performance requirement is designed and is obtained, this technology being known in the art.
Further, when free stream Mach number is more than MaminWhen, according to engine performance demand and the tune of Design of Inlet
Rule is saved, the hinge control adjustable type face rotation is adjusted, until free stream Mach number is Mamax。
Pass through the above-mentioned course of work so that air intake duct is in Mamin~MamaxIn the range of can reliable efficient operation, for burning
Room provides the compressed air stream for meeting the requirements such as speed, pressure, it is ensured that engine is in Mamin~MamaxIn the range of can work normally
And generate thrust performance.
Further, during rotation is adjusted, the work in the contract section adjustment type face 7 and expansion segment adjustment type face 8
Vertical range of the moved end apart from lip type face 5 remains equal.
Further, after the contract section adjustment type face 7 and expansion segment adjustment type face 8 rotate, the contract section tune
Nodal pattern face 7, contract segment type face 2, venturi type face 3, expansion segment type face 4 and expansion segment adjustment type face 8 surround a cavity, recessed
Intracavitary forms pneumatic venturi type face of the vortex as runner, to reduce inlet throat area, increases air intake duct shrinkage ratio, and then
Increase compression to incoming air, the corresponding venturi in the pneumatic venturi type face is highly defined as air intake duct larynx after the rotation of adjustable type face
Road height Hth, obtained by formula (1)：
Wherein, HthRunner venturi height after being rotated for adjustable type face, HcHeight is captured for air intake duct runner, q (Ma) is stream
Flow function, Ma0And MathRespectively free stream Mach number and runner venturi Mach number,For discharge coefficient, σthFor runner venturi stagnation pressure
Recovery coefficient.
Further, in low mach design conditions, such as Mamin, the contract section adjustment type face 7 and expansion segment
Contract segment type face 2 and expansion segment type face 4 are close in adjustment type face 8 respectively, with the increase of free stream Mach number, then need according to next
The parameters such as Mach number, the required inlet throat Mach number of Flow coefficient of inlet and combustion chamber are flowed, are calculated by above-mentioned formula (1)
The inlet throat height adjusted needed for obtaining, then be converted to contract section adjustment type face 7 and expansion segment adjusting by geometrical relationship
The angle rotated needed for type face 8, and free stream Mach number is bigger, 8 rotation angle of contract section adjustment type face 7 and expansion segment adjustment type face
Degree is bigger, and after forming pneumatic venturi, inlet throat circulation area is smaller, stronger to incoming compression, is provided for combustion chamber
The compressed air met the requirements, conducive to burning and engine performance raising is stablized.It in addition, when necessary can be by increasing contract section
Adjustment type face 7 and the rotation angle in expansion segment adjustment type face 8, completely close runner.
Above-mentioned rotation angle includes the rotation angle θ in contract section adjustment type face 77With the rotation in expansion segment adjustment type face 8
Angle, θ8, obtained by formula (2)；
Wherein, H7For 7 fixing end of contract section adjustment type face and 5 vertical height of lip type face, H8For expansion segment adjustment type face
8 fixing ends and 5 vertical height of lip type face, HthRunner venturi height after being rotated for adjustable type face, L7For contract section adjustment type
Distance between 7 two-end-point of face, L8The distance between 8 two-end-point of expansion segment adjustment type face, θ2For 2 two end point connecting line of contract segment type face with
Angle (taking acute angle) between venturi type face 3, θ4Angle between 4 two end point connecting line of expansion segment type face and venturi type face 3 (takes
Acute angle).
The rotation angle θ7And θ8When being 0 °, the contract section adjustment type face 7 and expansion segment adjustment type face 8 respectively with
Contract segment type face 2 and expansion segment type face 4 are bonded；
The rotation angle rotation angle θ7And θ8During for maximum value, the contract section adjustment type face 7 and expansion segment are adjusted
Type face 8 completely closes the runner, and the state is depending on special duty demand.
Further, the design of rectangle parallel way can be used in the subsonic combustion room and supersonic speed combustion chamber.
The design principle of the present invention is：
It is arranged in parallel, shared into row with the aloof punching engine in double combustion chamber and turbogenerator in view of combined engine
Gas system, while need to need to use non-with lifting body or Waverider aircraft integrated design, the aloof punching engine in double combustion chamber
Axial symmetry configuration, therefore for the nonaxisymmetrical configuration in double combustion chamber, the present invention separately designs corresponding air inlet runner, and is directed to
Each runner is designed, and is increased contract section adjustment type face and expansion segment adjustment type face, and pass through hinge and control its rotation, is being revolved
During turning, contract section adjustment type face, contract segment type face, venturi type face, expansion segment type face and expansion segment adjustment type face
A cavity can be surrounded, vortex is formed in cavity using as the pneumatic venturi type face of air inlet runner, by being received in rational design
The length in contracting segment type face, expansion segment type face and venturi type face, and respectively so that contract section adjustment type face and contract segment type
Face is consistent, and expansion segment adjustment type face is consistent with expansion segment type face length, so that contract section adjustment type face movable end and expansion
The distance for opening section adjustment type face movable end along flow direction is less than the characteristic dimension of vortex, and the characteristic dimension is obtained by CFD calculating,
And then stable vortex is formed in cavity；In addition, according to the size of rotation angle, different pneumatic venturi type faces will be obtained,
And then reduce inlet throat area in this way, increase air intake duct shrinkage ratio, that is, increase the compression to incoming air,
Inlet throat circulation area is smaller, stronger to incoming compression, provides the compressed air met the requirements for combustion chamber, is conducive to
Stablize burning and engine performance improves.Pass through above-mentioned adjusting so that air intake duct provides the air-flow for entering subsonic combustion room
Larger compression enters the tissue burning of subsonic combustion room, to entering bimodal ultrasound quick burning after a series of wave systems with subsonic speed
The air-flow for burning room provides smaller compression, and air-flow enters supersonic speed combustion chamber with supersonic speed and realizes that wide scope bimodal is burnt, with this
It widens engine work range of Mach numbers, improve engine performance.
Embodiment 2
The present embodiment provides a kind of above-mentioned air intake control method of double combustion chamber's Scramjet Inlet, including：
When flight Mach number reaches the first Mach number, the contract section adjustment type face of the sub- combustion runner and super burn runner
It is close to contract segment type face and expansion segment type face respectively with expansion segment adjustment type face, subsonic combustion room and supersonic speed combustion chamber are opened
Initial point fire, maintains gross thrust to meet aircraft demand, and the subsonic combustion room and supersonic speed combustion chamber are sub- combustion mode；
When flight Mach number reaches the second Mach number, the contract section adjustment type face of the sub- combustion runner and super burn runner and
Expansion segment adjustment type face proceeds by rotation, with the increase of flight Mach number, controls the interior receipts of sub- combustion runner and super burn runner
Contracting section adjustment type face and the rotation of expansion segment adjustment type face so that the subsonic combustion room and supersonic speed combustion chamber are in sub- combustion
Mode；
When flight Mach number reaches third Mach number, the contract section adjustment type face of super burn runner and expansion segment adjustment type face
It stops rotating adjusting, with the increase of flight Mach number, supersonic speed combustion chamber gradually changes from Asia combustion mode to super burn mode；
When flight Mach number reaches the 4th Mach number, supersonic speed combustion chamber is in super burn mode, completes mode transformation, with
The increase of flight Mach number, the contract section adjustment type face and expansion segment adjustment type face of super burn runner still stop, and do not rotate adjusting；
When flight Mach number reaches the 5th Mach number, the contract section adjustment type face of super burn runner and expansion segment adjustment type face
Start rotation to adjust, with the increase of flight Mach number so that the supersonic speed combustion chamber is in super burn mode always；
When flight Mach number reaches the 6th Mach number, this is engine maximum functional Mach number, and the Asia is fired runner and surpassed
The contract section adjustment type face and expansion segment adjustment type face for firing runner are stopped rotating adjusting, and shown subsonic combustion room is located always
Mode is fired in Asia, supersonic speed combustion chamber is in super burn mode always；
Between second Mach number and the 6th Mach number, contract section adjustment type face and the expansion segment of sub- combustion runner are controlled
Adjustment type face rotates so that the subsonic combustion room is always in sub- combustion mode.
It is described so that the subsonic combustion room and supersonic speed combustion chamber refer in sub- combustion mode so that the Asia always
Runner and the corresponding postrotational venturi Mach number of super burn runner is fired to be in 1.2~1.5 ranges always；
It is described so that the supersonic speed combustion chamber refers to always in super burn mode so that the corresponding rotation of the super burn runner
Venturi Mach number after turning is in 9/20~11/20 range of flight Mach number always；
The first Mach 2 ship Mamin；For punching engine start-up operation Mach number, the venturi horse of corresponding runner
Conspicuous number is Ma-th1；
The second Mach 2 ship Ma-2 corresponds to venturi Mach number Ma-th2, the Ma-th2 and is more than threshold value MaΔ1, institute
State MaΔ1 preferred scope is 1.2~1.5；
The third Mach 2 ship Ma-3, the 4th Mach 2 ship Ma-4, Ma-3 and Ma-4 and between realize supersonic speed
Transformation of the combustion chamber from sub- combustion mode to super burn mode, and the third Mach number is according to its corresponding venturi Mach number Ma-
Th3, the 4th Mach number Ma-4 and its correspond to venturi Mach number Ma-th4 and determine；Specifically, according to the 4th Mach number Ma-4 and
Venturi Mach number Ma-th4 can obtain rotation angle by formula (1) and (2), and Ma- can be obtained by rotation angle and Ma-th3
3；
The 4th Mach number Ma-4 preferred scopes are 5.5-6, and corresponding venturi Mach number Ma-th4 is 9/20~11/
20Ma-4；
The 5th Mach 2 ship Ma-5 corresponds to venturi Mach number Ma-th5, the Ma-th5 and is more than threshold value MaΔ2, institute
State MaΔ2 be 11/20Ma-5；
The 6th Mach 2 ship Mamax, punching engine maximum functional Mach number.
Unspecified part of the present invention is known to the skilled person technology.
Claims (10)
1. a kind of double combustion chamber's Scramjet Inlet, it is characterised in that：The air intake duct uses two dimensional inlet structure
Type, and sub- combustion runner and super burn runner are divided by support plate along flow direction, the sub- combustion channel is at least one, the super burn runner
At least two and divide equally in Asia combustion runner both sides；
The sub- combustion is consistent with super burn flow passage structure, is made of fixed face, adjustable type face and connects hinge, wherein, it is described solid
Sizing face includes external pressure miniature face, contract segment type face, venturi type face, expansion segment type face, lip type face and connects with combustion chamber
Direct type face；The adjustable type face includes contract section adjustment type face and expansion segment adjustment type face, the contract section adjustment type face
Contract section starting point and expansion segment endpoint are individually fixed in the one end in expansion segment adjustment type face, the other end is respectively movable end,
The connects hinge includes hinge a and b, and the hinge a and b is located at contract section starting point and expansion segment endpoint, and described respectively
Contract section adjustment type face and expansion segment adjustment type face can be rotated respectively around the hinge a and b by design rotation angle.
A kind of 2. double combustion chamber's Scramjet Inlet according to claim 1, which is characterized in that the interior receipts
Behind contracting section adjustment type face and the rotation of expansion segment adjustment type face, the contract section adjustment type face, contract segment type face, venturi type
Face, expansion segment type face and expansion segment adjustment type face surround a cavity, and gas of the vortex as air intake duct runner is formed in cavity
Dynamic venturi type face, the corresponding venturi in the pneumatic venturi type face are highly defined as runner venturi height H after the rotation of adjustable type faceth, lead to
Formula (1) is crossed to obtain：
Wherein, HthRunner venturi height after being rotated for adjustable type face, HcHeight is captured for air intake duct runner, q (Ma) is flow letter
Number, Ma0And MathRespectively free stream Mach number and runner venturi Mach number,For discharge coefficient, σthFor runner venturi total pressure recovery
Coefficient.
3. a kind of double combustion chamber's Scramjet Inlet according to claim 2, which is characterized in that in the hinge
During chain rotation is adjusted, movable end the hanging down apart from lip type face in the contract section adjustment type face and expansion segment adjustment type face
Straight distance remains equal.
A kind of 4. double combustion chamber's Scramjet Inlet according to claim 3, which is characterized in that the rotation
Angle includes the rotation angle θ in contract section adjustment type face7With the rotation angle θ in expansion segment adjustment type face8, obtained by formula (2)
It arrives；
Wherein, H7For contract section adjustment type face fixing end and lip type face vertical height, H8For expansion segment adjustment type face fixing end
With lip type face vertical height, HthRunner venturi height after being rotated for adjustable type face, L7For contract section adjustment type face two-end-point
Between distance, L8The distance between the two-end-point of expansion segment adjustment type face, θ2Between contract segment type face two end point connecting line and venturi type face
Sharp angle, θ4For the sharp angle between expansion segment type face two end point connecting line and venturi type face.
A kind of 5. double combustion chamber's Scramjet Inlet according to claim 4, which is characterized in that the rotation
Angle, θ7And θ8When being 0 °, the contract section adjustment type face and expansion segment adjustment type face respectively with contract segment type face and expansion
Segment type face paste is closed；The rotation angle θ7And θ8During for maximum value, the contract section adjustment type face and expansion segment adjustment type face with
It closes subject to the runner.
A kind of 6. double combustion chamber's Scramjet Inlet according to claim 5, which is characterized in that the sub- combustion
Runner is B runners, and the super burn runner is respectively A, C runner, and runner layout is selected from ABC, AABAA, ABABC, ABBC group
Any one of conjunction form, wherein A runners are identical with C runners.
7. according to a kind of double combustion chamber's Scramjet Inlet air inlet controlling party of claim 1-6 any one of them
Method, which is characterized in that including procedure below：
When flight Mach number reaches the first Mach number, contract section adjustment type face and the expansion of the sub- combustion runner and super burn runner
It opens section adjustment type face and is close to contract segment type face and expansion segment type face, subsonic combustion room and supersonic speed combustion chamber starting point respectively
Fire, the subsonic combustion room and supersonic speed combustion chamber are sub- combustion mode；
When flight Mach number reaches the second Mach number, contract section adjustment type face and the expansion of the sub- combustion runner and super burn runner
Section adjustment type face proceeds by rotation, with the increase of flight Mach number, controls the contract section of sub- combustion runner and super burn runner
Adjustment type face and the rotation of expansion segment adjustment type face so that the subsonic combustion room and supersonic speed combustion chamber are in sub- combustion mould
State；
When flight Mach number reaches third Mach number, the contract section adjustment type face of super burn runner and expansion segment adjustment type face stop
Rotation is adjusted, and with the increase of flight Mach number, supersonic speed combustion chamber gradually changes from Asia combustion mode to super burn mode；
When flight Mach number reaches the 4th Mach number, supersonic speed combustion chamber is in super burn mode, mode transformation is completed, with flight
The increase of Mach number, the contract section adjustment type face and expansion segment adjustment type face of super burn runner still stop, and do not rotate adjusting；
When flight Mach number reaches the 5th Mach number, the contract section adjustment type face and expansion segment adjustment type face of super burn runner start
Rotation is adjusted, with the increase of flight Mach number so that the supersonic speed combustion chamber is in super burn mode always；
When flight Mach number reaches the 6th Mach number, contract section adjustment type face and the expansion of the sub- combustion runner and super burn runner
Section adjustment type face is stopped rotating adjustings, and mode is fired in shown subsonic combustion room in sub- always, and supersonic speed combustion chamber is located always
In super burn mode；
Between second Mach number and the 6th Mach number, the contract section adjustment type face and expansion segment that control sub- combustion runner are adjusted
Type face rotates so that the subsonic combustion room is always in sub- combustion mode.
8. the method according to the description of claim 7 is characterized in that further include at least one of following technical characteristic：
The first Mach 2 ship Mamin；The venturi Mach 2 ship Ma-th1 of its corresponding runner；
The second Mach 2 ship Ma-2 corresponds to venturi Mach number Ma-th2, the Ma-th2 and is more than threshold value MaΔ1, it is described
MaΔ1 ranging from 1.2~1.5；
The third Mach 2 ship Ma-3, the 4th Mach 2 ship Ma-4, Ma-3 and Ma-4 and between realize supersonic combustion
Transformation of the room from sub- combustion mode to super burn mode, and the third Mach number is according to its corresponding venturi Mach number Ma-th3, the
Four Mach number Ma-4 correspond to venturi Mach number Ma-th4 to determine with it；
The 4th Mach number Ma-4 ranging from 5.5-6, corresponding venturi Mach number Ma-th4 are 9/20~11/20Ma-4；
The 5th Mach 2 ship Ma-5 corresponds to venturi Mach number Ma-th5, the Ma-th5 and is more than threshold value MaΔ2, it is described
MaΔ2 be 11/20Ma-5；
The 6th Mach 2 ship Mamax, engine maximum functional Mach number.
It is 9. according to the method described in claim 8, it is characterized in that, described so that the subsonic combustion room is always in sub- combustion
Mode refers to so that the corresponding postrotational venturi Mach number of the sub- combustion runner is in 1.1~1.5 ranges always.
It is 10. according to the method described in claim 9, it is characterized in that, described so that the supersonic speed combustion chamber is always in super
Combustion mode refer to so that the corresponding postrotational venturi Mach number of the super burn runner always in flight Mach number 9/20~
In 11/20 range.
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| CN107605601B (en) * | 2017-09-29 | 2018-08-31 | 南京航空航天大学 | A kind of Supersonic Inlet and its control method and design method of capture area and throat area synchronous adjustment |
| CN108317541B (en) * | 2018-02-26 | 2020-07-07 | 中国科学院力学研究所 | Ramjet engine |
| CN108592085B (en) * | 2018-03-09 | 2020-10-16 | 西北工业大学 | Variable geometry supersonic combustion chamber |
| CN108825381B (en) * | 2018-06-22 | 2019-06-07 | 西北工业大学 | A kind of rocket based combined cycle engine surface compression change geometry air intake duct |
| CN109098859B (en) * | 2018-08-13 | 2019-10-18 | 厦门大学 | A kind of hypersonic interior Parallet three-way valve road intake valve deposit mechanism design method |
| CN109667670B (en) * | 2019-01-24 | 2021-03-09 | 南京航空航天大学 | Vortex-controlled adjustable valve device for assisting start of ultra/hypersonic binary air inlet |
| CN114427496B (en) * | 2020-10-29 | 2023-07-25 | 南京理工大学 | Variable-geometry supersonic air inlet device with leakage channel |
| CN113700560B (en) * | 2021-09-24 | 2023-01-20 | 西安航天动力研究所 | Half-film overturning supersonic speed adjustable air inlet channel |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5165227A (en) * | 1990-04-14 | 1992-11-24 | Mtu Motoren- Und Turbinen-Union Munchen Gmbh | Propelling nozzle for a hypersonic engine |
| US7762077B2 (en) * | 2006-12-05 | 2010-07-27 | Pratt & Whitney Rocketdyne, Inc. | Single-stage hypersonic vehicle featuring advanced swirl combustion |
| CN103790735A (en) * | 2014-01-24 | 2014-05-14 | 西北工业大学 | Variable-structure exhaust nozzle of rocket based combined cycle engine |
| CN104110326A (en) * | 2014-07-02 | 2014-10-22 | 北京航空航天大学 | New concept high-speed aerocraft propulsion system layout method |
| CN205578133U (en) * | 2016-03-29 | 2016-09-14 | 杨庆春 | Super combustion pressed engine |
| CN106150757A (en) * | 2016-08-10 | 2016-11-23 | 西北工业大学 | A kind of dual pathways becomes geometry rocket based combined cycle electromotor |
-
2017
- 2017-02-17 CN CN201710086436.0A patent/CN107013334B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5165227A (en) * | 1990-04-14 | 1992-11-24 | Mtu Motoren- Und Turbinen-Union Munchen Gmbh | Propelling nozzle for a hypersonic engine |
| US7762077B2 (en) * | 2006-12-05 | 2010-07-27 | Pratt & Whitney Rocketdyne, Inc. | Single-stage hypersonic vehicle featuring advanced swirl combustion |
| CN103790735A (en) * | 2014-01-24 | 2014-05-14 | 西北工业大学 | Variable-structure exhaust nozzle of rocket based combined cycle engine |
| CN104110326A (en) * | 2014-07-02 | 2014-10-22 | 北京航空航天大学 | New concept high-speed aerocraft propulsion system layout method |
| CN205578133U (en) * | 2016-03-29 | 2016-09-14 | 杨庆春 | Super combustion pressed engine |
| CN106150757A (en) * | 2016-08-10 | 2016-11-23 | 西北工业大学 | A kind of dual pathways becomes geometry rocket based combined cycle electromotor |
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