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A new design of a support force measuring system for hypersonic vehicle aerodynamic measurement
Flow Measurement and Instrumentation 70 (2019) 101646
Contents lists available at ScienceDirect
Flow Measurement and Instrumentation journal homepage: http://www.elsevier.com/locate/flowmeasinst
Short communication
A new design of a support force measuring system for hypersonic vehicle aerodynamic measurement Bokai Liu, Hongli Gao *, Kang Zhao, Juting Huang, Yi Sun School of Mechanical Engineering, Southwest Jiaotong University, Chengdu 610031, PR China
A R T I C L E I N F O
A B S T R A C T
Keywords: Force measurement Hypersonic vehicle Wind tunnel test Balance Sensor
In order to simulate the flight state of the aerocraft better, the aerodynamic parameters are obtained by inte grating the airframe and propulsion system in a highly integrated configuration and ignition of the engine during the wind tunnel test. At present, the commonly used internal balance scheme forms a system composed of the model, internal balance and the support. However, most hypersonic vehicle models are flat or slender, which makes it difficult to provide installation space for internal balance, improves the difficulty of model design. This note proposes a support force measuring system. By integrating the balance into the structure of the support, it no longer occupies functional space in the cavity of the model, and its special structure can ensure that the addi tional torque generated by thrust/drag cannot act on the torque measuring element. This study provides a new way of thinking for the integrated model test with propulsion system.
1. Introduction In the past wind tunnel tests, the aerodynamic characteristics of the vehicle, such as thrust/drag, lift and pitch moment, were obtained by force measurement tests in wind tunnel with the structure of the test model-internal balance-support and the internal balance of side wall support, abdomen support or back support [1–3]. However, in the thrust drag measurement test of hypersonic vehicle with engines, as that structure and aerodynamic configuration of the hypersonic vehicle are highly integrated, the interior of the airframe is mostly functionally designed for arrange components such as inlet ports, engines, fuel in jection block, oil pipelines and the like [4]. The method for separating the engine from the test model is no longer applicable, and the hyper sonic vehicle is mostly slender or flat [5,6]. Therefore, it is difficult to provide space for the installation of the internal balance, which greatly increases the difficulty of the aircraft model and the internal balance designer. At present, some methods can be used to solve the problem of in ternal balance occupying model space and measurement error. Klaus Hannemann, Jan Martinez Schramm [7,8] et al. designed a cut-off ten sion line for LAPCAT II ramjet thrust/drag measurement, and aero dynamic characteristics of the model are obtained by solving the relationship between displacement and time. Dufrene [9] et al. designed an elastic connection and high speed camera with six-component
acceleration balance scheme and obtained the aerodynamic data of the test model in a test. Bokai Liu [10] et al. carried out a force measurement method of suspended test model. This scheme solved the problem of the balance occupying the internal functional space of the model. TANNO Hideyuki [11] et al. revealed the potential of the free-light aerodynamic test method in shock tunnel force testing by means of electromagnetic suspension with free release and acceleration sensors acquisition during testing. Tavakolpour-Saleh A R [12] designed a new type of three components external strain balance. The balance was installed on the back of the model without occupying the internal space of the model. Vadassery [13] developed an external stress wave balance. The dynamic performance of the balance and the position strain of the patch were analyzed. M.L.C.C. Reis [14] carried out a study on the evaluation of calibration uncertainty for a six-components external balance. Almeida [15] developed a ring-type balance to measure drag, lift and pitching moment in wind tunnel tests. Shixiong Zhang [16] carried out a me chanical analysis of the disturbance of the axial force and proposed a method to modify the Whiston bridge equation. J. van der Vooren [17] discussed the influence of compressible flow on aerodynamics of the model, and the correction of buoyancy is also discussed in consideration of thrust. At present, no structural design scheme based on the principle of spatial concurrent force system has been found in the literature to realize the external shift of the internal balance and the elimination of interference.
* Corresponding author. E-mail address: [email protected] (H. Gao). https://doi.org/10.1016/j.flowmeasinst.2019.101646 Received 1 July 2019; Received in revised form 20 August 2019; Accepted 30 September 2019 Available online 2 October 2019 0955-5986/© 2019 Elsevier Ltd. All rights reserved.
B. Liu et al.
Flow Measurement and Instrumentation 70 (2019) 101646
In this note, a new type of support force measuring system is pro posed. The force measuring system changed the structure of the force measuring system composed of the traditional aircraft model-internal balance-support. The internal balance and the abdominal-support are integrated, and the strain gauges are placed on the surface of the load measuring structure of the support, formed a force-measuring system composed of the aircraft model and the load measuring support. The load measuring support is fixedly connected with the abdomen of the aircraft model directly, and does not occupy the functional space inside the model frame, thus greatly reducing the design difficulty of the in tegrated aircraft model with engine. Different from the traditional external strain balance, the moment reference center of the system is still inside the test model and coincides with the moment reference center of the model. When the model is subjected to thrust/drag force Fx and lift force Fy, additional moment will not act on moment measuring element, which effectively limits the interference of forces on moment measure ment. As that force balance and the support are integrated into one structure, the whole force measuring system also has greater stiffness and better linearity.
Fig. 2. Structure and bridge circuit layout of support force measuring system.
When the aerodynamic load acting on the test model is transmitted to the load measuring support, the pitching moment Mz makes the inner small triangular structure of the moment measuring part rotate around the transverse axis (Z axis) of the moment reference center of the sup port, and the moment measuring beam becomes the cantilever beam structure to be effectively deformed; To the thrust/drag force Fx, when it passes through the intersection point of the extension line supported by the two force bars, and the bearing reaction force and thrust/drag force Fx of the two force bars support form a balanced spatial concurrent force system through the moment reference center. Therefore, large triangle structure of the upper part of the support rotates integrally, at the same time the small triangle structure of the moment measuring part does not rotate. So the moment measuring beam does not deform, hence one can see that the force does not interfere with the moment. Then the decomposition of the force and the moment subjected to the aero dynamic force is realized. According to the layout of the Wheatstone Bridge, the supply bridge voltage is set to U, and the loaded output voltages of the three components are expressed by the following equations:
2. Principles of measurement The support force measuring system is used to measure thrust/drag force, lift force and pitching moment in hypersonic wind tunnel force measuring test. According to the computational fluid dynamics (CFD) experiments results of the aerocraft test model, the maximum design loads of thrust/drag force, lift force and pitching moment are 2000 N, 3000 N and 1000Nm respectively. Brief description of the drawings Fig. 1 is an illustration of the component parts of the force measuring support, consisted of a main support and front and rear fairings. The upper connecting plate of the support is fixedly connected with the test model, the hole position is arranged according to the specific model, and the lower part of the support is fixedly connected with the ground of the wind tunnel test section. Fig. 2 shows that structure of the force measuring support and the design form of the bridge circuit. The moment measurement part of the system is composed of a pair of oblique hinge double-bar support and a patch sensitive beam structure, and the moment reference center of the support dynamometer system is located at the intersection point of the extension line of the pair of oblique hinge double-bar support and also intersects with the center line of the pitch moment patch beam. It is necessary to ensure that the torque reference center coincides with the torque reference center of the test model, which can be achieved by adjusting the angle and position of the doublebar support.
ΔUFx ¼
R11R13 R12R14 ⋅U ðR11 þ R12ÞðR13 þ R14Þ
(1)
ΔUFy ¼
ðR21 þ R22ÞðR25 þ R26Þ ðR23 þ R24ÞðR27 þ R28Þ ⋅U ðR21 þ R22 þ R23 þ R24ÞðR25 þ R26 þ R27 þ R28Þ
(2)
R31R33 R32R34 ⋅U ðR31 þ R32ÞðR33 þ R34Þ
(3)
ΔUMz ¼
3. Simulation calculation and analysis In an impulse combustion wind tunnel, the aircraft model will be subjected to an impact load several times the design load [18]. In order to protect the measuring circuit, it is necessary to control the strain of each strain gauge application surface to less than 150 micro-strains, so strain analysis is required. ANSYS was used to conduct mechanical simulation, and the mesh control is applied to the strain gauges surface. The hexahedron element is adopted, and the element size is controlled to 3 mm. A total of 109824 elements were generated, and the grid inde pendence is verified after calculation. Figs. 3–5 show the strain state of the aircraft model under the full-scale loads of thrust drag, lift and pitch moment separately. As can be seen from Fig. 3, the drag measurement sensitive structure produces about 114 microstrains in the effective strain direction of the strain gauge (parallel to the Y axis) when the support dynamometer system is loaded with a full-scale load in the thrust drag direction. The interference of drag load on lift measurement and pitch moment mea surement is eliminated by electrical method and mechanical method. The electrical method reduces the interference between components through the reasonable layout of the Whiston Bridge and the mechanical method can avoid or reduce interference by optimizing the mechanical
Fig. 1. Composition of support force measuring system. 2
B. Liu et al.
Flow Measurement and Instrumentation 70 (2019) 101646
Fig. 3. Thrust and drag load strain results. Fig. 5. Pitching moment load strain results.
changed theoretically, so the interference of thrust and drag loading on lift measurement is eliminated by electric method. As shown in the figure, the effective direction (parallel direction along the X axis) strain is 31.4 micro-strains and 31.3 micro-strains in No.21 and No.22 gauge position, respectively. The total resistance of the bridge arm formed by two strain gauges of the same type tends to zero after forming one bridge arm at the above two positions, which will have a good effect of elimi nating disturbance after post calibrating. For the interference of drag on pitching moment, as shown in Fig. 2, the thrust force can only make the force-moment decomposition mechanism of large triangle rotate as a whole. Theoretically, the torque measuring element does not deform. As shown in the simulation results of Fig. 3, only about 1.4 micro-strains are produced in the effective strain direction (along the Y-axis parallel di rection) of the pitch moment element patch position when the drag is loaded, which achieves a good effect of eliminating the interference. The main reason of the strain is that the moment reference center of the support is slightly shifted after the force is applied. As seen in Fig. 4, that lift measurement sensitive structure produce about 88 micro-strains in the effective strain direction of the strain gauge (parallel direction along the X-axis) when the support force measurement system is loaded with a full-scale load in the lift direction. The lift force has little influence on the thrust or drag measurement, and the effective direction of strain gauge application produces about 3 strain quantities of micro-strain magnitude. For the disturbance of lift to pitching moment, as that lift load position passes through the intersec tion point of the spatial concurrent force system, the lift force only cause the tension and compression deformation of the pitch moment measuring beam, and the strain at the patch position is in the same di rection and the absolute value is the same, so the pitch bridge theory does not cause unbalance, the output voltage does not increase, and the interference of the lift force on the pitch moment measuring is eliminated. As seen in Fig. 5, that pitch moment measurement sensitive structure produce about 93 microstrains in the effective strain direction (parallel direction along the Y axis) of the strain gauge when the support force measurement system is loaded with a full-scale load in the pitch moment
Fig. 4. Lift force load strain results.
structure of the measuring system. Lift measurement bridge adopts eight strain gauges combination bridge form. The two strain gauges positions of the same bridge arm are symmetrical about the Y axis centerline of the thrust or drag measuring element beam. When disturbance strain occurs, the two strain gauges resistances of the same bridge arm increase and decrease exactly one by one, and the absolute values are the same, which ensures that the resistance values of the four bridge arms are not 3
B. Liu et al.
Flow Measurement and Instrumentation 70 (2019) 101646
direction. The pitching moment has little influence on the thrust or drag measurement, and the effective direction of strain gauge application produces about 4.5 strain magnitudes. The interference of pitching moment to lift force is also eliminated by electrical method. Although the effective strain of each strain gauge is about 31 micro-strains, the strain directions of two strain gauges in the same bridge arm are opposite. Theoretically, the total resistance of each arm does not change, and there is no output voltage increment of the lift bridge. Table 1 shows the principal strain and the relative principal strain disturbance strain of three measurement directions under each compo nent individual full-scale loading. The strain is calculated as the strain of each bridge arm. The maximum disturbance is the disturbance of pitching moment loading to thrust or drag measurement, which is 3.9%. In impulse combustion wind tunnel test, the system will bear several times the design load of hypersonic airflow impact load [19], so the strength analysis must be carried out. The allowable stress is calculated according to the maximum possible transient load, and the allowable stress formula is shown in Equation (4).
σ � ½σ � ¼
σb
Fig. 6. Von-Mises stress cloud diagram.
(4)
nc ⋅na
Where σb is the yield stress of 00Ni18Co8Mo5TiAl, which is 1750 MPa nc and na represent the impact factor and safety factor of the wind tunnel, respectively. According to the test requirements, if the impact factor is 3 and the safety factor is 2, the design safety factor n ¼ nc � na ¼ 6. The maximum design stress σ should not be greater than the allowable stress [σ]. The allowable stress [σ] is 291.7 MPa. All components of the model are loaded synthetically, including the thrust/drag force is 2000 N, the lift force is 3000 N, and the pitching moment is 1000Nm. The calculated Von-Mises stress cloud diagram is shown in Fig. 6. It can be seen from the figure that the maximum stress concentration point appears on the pitching moment measuring beam structure under comprehensive loading. The stress value is 194 MPa, which is less than the allowable stress 291.7 MPa. The system meets the test requirements in strength. Because the effective force measurement time of the impulse com bustion wind tunnel used by the force measurement system is very short, the average force measurement time is only 250 ms. In order to ensure the accuracy of the force measurement results, at least 6 cycles of data shall be obtained during the test period [10,20]. According to the test requirements, the first natural frequency of the system shall reach 30 Hz. Therefore, modal analysis is carried out on the simplified structure of the model and the assembly of the support force measuring system, the model material is aviation hard aluminum 7075, and the total mass is about 200 Kg. Fig. 7 shows the first six modal modes of the force measuring system under the constraint condition at the bottom of the support. Table 2 shows the calculation results of the first six modal natural frequencies. The results show that the first order natural fre quency of the force measuring system reaches 37 Hz. The first vibration of the thrust drag, lift and pitching moment measured by the dyna mometer system occurs at the fourth order 204 Hz, the sixth order 298 Hz and the third order 78 Hz, respectively, which meets the test requirements.
Fig. 7. The first six modal modes. Table 2 First six natural frequencies and main vibration directions.
4. Conclusion In summary, we proposed a new type of hypersonic wind tunnel
Modal order
Frequency (Hz)
Main Vibration Direction
1 2 3 4 5 6
37.6 38.7 78.8 204.6 285.6 298.6
Lateral direction Yaw direction Pitching direction Thrust/Drag direction Alignment direction Lift direction
Table 1 Main strain and interference strain loaded by each component separately.
Thrust or drag Fx ¼ 2000 N Lift Fy ¼ 3000 N Pitching moment Mz ¼ 1000Nm
Thrust or drag strain
Interference with thrust or drag
Lift strain
Interference with lift
Pitching moment strain
Interference with pitching moment
�114με �3.1με �4.5με
\ �2.7% �3.9%
＜0.5με �88με ＜1.5με
0.6%＜ \ 1.7%＜
�1.4με �0 �93με
�1.5% �0% \
4
B. Liu et al.
Flow Measurement and Instrumentation 70 (2019) 101646
org/10.1016/j.flowmeasinst.2019.101646.
support force measuring system, simulated and analyzed the strain sensitivity, the interference between components and the modal attri butes of each measurement component of the support force measuring system. The results show that all the parameters meet the requirements of hypersonic wind tunnel load test. The design also provides a new idea for wind tunnel load test. The main advantages of the design are as follows:
References [1] A.L. Smith, D.J. Mee, W.J.T. Daniel, et al., Design, modelling and analysis of a six component force balance for hypervelocity wind tunnel testing[J], Comput. Struct. 79 (11) (2001) 1077–1088. [2] N.P. Joshi, Strain-gauge applications in wind tunnel balances[J], IETE Tech. Rev. 9 (2) (1992) 161–167. � ci�c, et al., Wind tunnel measurement of small [3] A. Boutemedjet, M. Samard�zi�c, D. Cur� values of rolling moment using six-component strain gauge balance, Measurement 116 (2018) 438–450. [4] W. Jifei, C.A.I. Jinsheng, L.I.U. Chuanzhen, et al., Aerodynamic configuration integration design of hypersonic cruise aircraft with inward-turning inlets[J], Chin. J. Aeronaut. 30 (4) (2017) 1349–1362. [5] V. Morio, F. Cazaurang, P. Vernis, Flatness-based hypersonic reentry guidance of a lifting-body vehicle[J], Contr. Eng. Pract. 17 (5) (2009) 588–596. [6] J. Liu, F. Ding, W. Huang, et al., Novel approach for designing a hypersonic gliding–cruising dual waverider vehicle[J], Acta Astronaut. 102 (2014) 81–88. [7] K. Hannemann, J. Martinez Schramm, S. Laurence, et al., Free flight testing of a scramjet engine in a large scale shock tunnel[C]. 20th, in: AIAA International Space Planes and Hypersonic Systems and Technologies Conference, AIAA, Glasgow, Scotland, 2015. [8] J.M. Schramm, K. Hannemann, S. Karl, et al., Ground Testing Synthesis of the LAPCAT II Small Scale Flight Experiment Configuration Scramjet Flow Path[C], Int. Space Planes and Hypersonic Systems and Technology Conference, 2013. [9] A. Dufrene, M.G. MacLean, M.S. Holden, Missile Stage Separation Testing in CUBRC Hypersonic Tunnels[C], AIAA Missile Sciences Conference, Monterey, CA, November 2010. [10] B. Liu, S. Li, H. Gao, et al., Suspension Force Measuring System for Hypersonic Wind Tunnel Test: Design and tests[J]. Measurement, 2019. [11] H. Tanno, T. Komuro, K. Sato, et al., Free-flight aerodynamic test of elliptic cone in Shock Tunnel[C], in: //20th AIAA International Space Planes and Hypersonic Systems and Technologies Conference, 2015, p. 3655. [12] A.R. Tavakolpour-Saleh, A.R. Setoodeh, M. Gholamzadeh, A novel multicomponent strain-gauge external balance for wind tunnel tests: simulation and experiment[J], Sens. Actuators A Phys. 247 (2016) 172–186. [13] P. Vadassery, D.D. Joshi, T.C. Rolim, et al., Design and testing of an external drag balance for a hypersonic shock tunnel, Measurement 46 (7) (2013) 2110–2117. [14] M. Reis, R.M. Castro, O.A.F. Mello, Calibration uncertainty estimation of a straingage external balance, Measurement 46 (1) (2013) 24–33. [15] R.A.B. Almeida, D.C. Vaz, A.P.V. Urgueira, et al., Using ring strain sensors to measure dynamic forces in wind-tunnel testing, Sens. Actuators A Phys. 185 (2012) 44–52. [16] S. Zhang, X. Li, H. Ma, et al., Mechanical analysis of normal force interference on axial force measurement for internal sting balance, Aero. Sci. Technol. 58 (2016) 351–357. [17] J. Van der Vooren, On drag and lift analysis of transport aircraft from wind tunnel measurements[J], Aero. Sci. Technol. 12 (4) (2008) 337–345. [18] Q. Zhang, B. Qin, H. Yan, et al., A methodology to predict shock overpressure decay in a tunnel produced by a premixed methane/air explosion[J], J. Loss Prev. Process. Ind. 44 (2016) 275–281. [19] H.K. Rafsanjani, Assessment of image processing methods to characterize the precise shape of the jet head in the hypersonic jet flows[J], Flow Meas. Instrum. 64 (2018) 142–154. [20] D.L. Schultz, T.V. Jones, Heat-Transfer Measurements in Short-Duration Hypersonic Facilities, Agardograph, 1973, p. 165.
1. The support force measuring system changed the model-internal balance-support configuration in hypersonic wind tunnel test, the system structure composed of the model and the load-measuring support is formed, which does not need to design the internal bal ance separately and does not occupy the internal space of the test model, thus greatly facilitating the layout and design of the inte grated model with engines. The external position of the loadmeasuring structure also makes the installation and maintenance of the system more convenient. 2. Different from the traditional external balance, the special structure of the concurrent force system of the dynamometer system greatly reduces the disturbance of the additional moment caused by the thrust/drag and lift to the pitching moment. The finite element simulation results show that the additional moment produced by thrust force only makes the whole concurrent force system structure rotate, the pitching moment measuring element has almost no effective strain, the lift force only makes the tensile and compressive deformation of the pitching moment measuring beam, and the electrical decomposition of Wheatstone bridge makes the interfer ence between the measuring components very small. 3. The support force measuring system eliminates the bolt connection between the traditional balance and the support by integrating the traditional internal balance with the support. Therefore, the support force measuring system has greater stiffness, higher response fre quency and better linearity, and is more suitable for the instanta neous dynamic force measurement in hypersonic wind tunnel test. Acknowledge This work is financially supported by the National Natural Science Foundation of China (51775452 and 51805457) and the Fundamental Research Funds for Central Universities of China (2682019CX35, 2682017ZDPY09 and 2018GF02). Appendix A. Supplementary data Supplementary data to this article can be found online at https://doi.
5
Contents lists available at ScienceDirect
Flow Measurement and Instrumentation journal homepage: http://www.elsevier.com/locate/flowmeasinst
Short communication
A new design of a support force measuring system for hypersonic vehicle aerodynamic measurement Bokai Liu, Hongli Gao *, Kang Zhao, Juting Huang, Yi Sun School of Mechanical Engineering, Southwest Jiaotong University, Chengdu 610031, PR China
A R T I C L E I N F O
A B S T R A C T
Keywords: Force measurement Hypersonic vehicle Wind tunnel test Balance Sensor
In order to simulate the flight state of the aerocraft better, the aerodynamic parameters are obtained by inte grating the airframe and propulsion system in a highly integrated configuration and ignition of the engine during the wind tunnel test. At present, the commonly used internal balance scheme forms a system composed of the model, internal balance and the support. However, most hypersonic vehicle models are flat or slender, which makes it difficult to provide installation space for internal balance, improves the difficulty of model design. This note proposes a support force measuring system. By integrating the balance into the structure of the support, it no longer occupies functional space in the cavity of the model, and its special structure can ensure that the addi tional torque generated by thrust/drag cannot act on the torque measuring element. This study provides a new way of thinking for the integrated model test with propulsion system.
1. Introduction In the past wind tunnel tests, the aerodynamic characteristics of the vehicle, such as thrust/drag, lift and pitch moment, were obtained by force measurement tests in wind tunnel with the structure of the test model-internal balance-support and the internal balance of side wall support, abdomen support or back support [1–3]. However, in the thrust drag measurement test of hypersonic vehicle with engines, as that structure and aerodynamic configuration of the hypersonic vehicle are highly integrated, the interior of the airframe is mostly functionally designed for arrange components such as inlet ports, engines, fuel in jection block, oil pipelines and the like [4]. The method for separating the engine from the test model is no longer applicable, and the hyper sonic vehicle is mostly slender or flat [5,6]. Therefore, it is difficult to provide space for the installation of the internal balance, which greatly increases the difficulty of the aircraft model and the internal balance designer. At present, some methods can be used to solve the problem of in ternal balance occupying model space and measurement error. Klaus Hannemann, Jan Martinez Schramm [7,8] et al. designed a cut-off ten sion line for LAPCAT II ramjet thrust/drag measurement, and aero dynamic characteristics of the model are obtained by solving the relationship between displacement and time. Dufrene [9] et al. designed an elastic connection and high speed camera with six-component
acceleration balance scheme and obtained the aerodynamic data of the test model in a test. Bokai Liu [10] et al. carried out a force measurement method of suspended test model. This scheme solved the problem of the balance occupying the internal functional space of the model. TANNO Hideyuki [11] et al. revealed the potential of the free-light aerodynamic test method in shock tunnel force testing by means of electromagnetic suspension with free release and acceleration sensors acquisition during testing. Tavakolpour-Saleh A R [12] designed a new type of three components external strain balance. The balance was installed on the back of the model without occupying the internal space of the model. Vadassery [13] developed an external stress wave balance. The dynamic performance of the balance and the position strain of the patch were analyzed. M.L.C.C. Reis [14] carried out a study on the evaluation of calibration uncertainty for a six-components external balance. Almeida [15] developed a ring-type balance to measure drag, lift and pitching moment in wind tunnel tests. Shixiong Zhang [16] carried out a me chanical analysis of the disturbance of the axial force and proposed a method to modify the Whiston bridge equation. J. van der Vooren [17] discussed the influence of compressible flow on aerodynamics of the model, and the correction of buoyancy is also discussed in consideration of thrust. At present, no structural design scheme based on the principle of spatial concurrent force system has been found in the literature to realize the external shift of the internal balance and the elimination of interference.
* Corresponding author. E-mail address: [email protected] (H. Gao). https://doi.org/10.1016/j.flowmeasinst.2019.101646 Received 1 July 2019; Received in revised form 20 August 2019; Accepted 30 September 2019 Available online 2 October 2019 0955-5986/© 2019 Elsevier Ltd. All rights reserved.
B. Liu et al.
Flow Measurement and Instrumentation 70 (2019) 101646
In this note, a new type of support force measuring system is pro posed. The force measuring system changed the structure of the force measuring system composed of the traditional aircraft model-internal balance-support. The internal balance and the abdominal-support are integrated, and the strain gauges are placed on the surface of the load measuring structure of the support, formed a force-measuring system composed of the aircraft model and the load measuring support. The load measuring support is fixedly connected with the abdomen of the aircraft model directly, and does not occupy the functional space inside the model frame, thus greatly reducing the design difficulty of the in tegrated aircraft model with engine. Different from the traditional external strain balance, the moment reference center of the system is still inside the test model and coincides with the moment reference center of the model. When the model is subjected to thrust/drag force Fx and lift force Fy, additional moment will not act on moment measuring element, which effectively limits the interference of forces on moment measure ment. As that force balance and the support are integrated into one structure, the whole force measuring system also has greater stiffness and better linearity.
Fig. 2. Structure and bridge circuit layout of support force measuring system.
When the aerodynamic load acting on the test model is transmitted to the load measuring support, the pitching moment Mz makes the inner small triangular structure of the moment measuring part rotate around the transverse axis (Z axis) of the moment reference center of the sup port, and the moment measuring beam becomes the cantilever beam structure to be effectively deformed; To the thrust/drag force Fx, when it passes through the intersection point of the extension line supported by the two force bars, and the bearing reaction force and thrust/drag force Fx of the two force bars support form a balanced spatial concurrent force system through the moment reference center. Therefore, large triangle structure of the upper part of the support rotates integrally, at the same time the small triangle structure of the moment measuring part does not rotate. So the moment measuring beam does not deform, hence one can see that the force does not interfere with the moment. Then the decomposition of the force and the moment subjected to the aero dynamic force is realized. According to the layout of the Wheatstone Bridge, the supply bridge voltage is set to U, and the loaded output voltages of the three components are expressed by the following equations:
2. Principles of measurement The support force measuring system is used to measure thrust/drag force, lift force and pitching moment in hypersonic wind tunnel force measuring test. According to the computational fluid dynamics (CFD) experiments results of the aerocraft test model, the maximum design loads of thrust/drag force, lift force and pitching moment are 2000 N, 3000 N and 1000Nm respectively. Brief description of the drawings Fig. 1 is an illustration of the component parts of the force measuring support, consisted of a main support and front and rear fairings. The upper connecting plate of the support is fixedly connected with the test model, the hole position is arranged according to the specific model, and the lower part of the support is fixedly connected with the ground of the wind tunnel test section. Fig. 2 shows that structure of the force measuring support and the design form of the bridge circuit. The moment measurement part of the system is composed of a pair of oblique hinge double-bar support and a patch sensitive beam structure, and the moment reference center of the support dynamometer system is located at the intersection point of the extension line of the pair of oblique hinge double-bar support and also intersects with the center line of the pitch moment patch beam. It is necessary to ensure that the torque reference center coincides with the torque reference center of the test model, which can be achieved by adjusting the angle and position of the doublebar support.
ΔUFx ¼
R11R13 R12R14 ⋅U ðR11 þ R12ÞðR13 þ R14Þ
(1)
ΔUFy ¼
ðR21 þ R22ÞðR25 þ R26Þ ðR23 þ R24ÞðR27 þ R28Þ ⋅U ðR21 þ R22 þ R23 þ R24ÞðR25 þ R26 þ R27 þ R28Þ
(2)
R31R33 R32R34 ⋅U ðR31 þ R32ÞðR33 þ R34Þ
(3)
ΔUMz ¼
3. Simulation calculation and analysis In an impulse combustion wind tunnel, the aircraft model will be subjected to an impact load several times the design load [18]. In order to protect the measuring circuit, it is necessary to control the strain of each strain gauge application surface to less than 150 micro-strains, so strain analysis is required. ANSYS was used to conduct mechanical simulation, and the mesh control is applied to the strain gauges surface. The hexahedron element is adopted, and the element size is controlled to 3 mm. A total of 109824 elements were generated, and the grid inde pendence is verified after calculation. Figs. 3–5 show the strain state of the aircraft model under the full-scale loads of thrust drag, lift and pitch moment separately. As can be seen from Fig. 3, the drag measurement sensitive structure produces about 114 microstrains in the effective strain direction of the strain gauge (parallel to the Y axis) when the support dynamometer system is loaded with a full-scale load in the thrust drag direction. The interference of drag load on lift measurement and pitch moment mea surement is eliminated by electrical method and mechanical method. The electrical method reduces the interference between components through the reasonable layout of the Whiston Bridge and the mechanical method can avoid or reduce interference by optimizing the mechanical
Fig. 1. Composition of support force measuring system. 2
B. Liu et al.
Flow Measurement and Instrumentation 70 (2019) 101646
Fig. 3. Thrust and drag load strain results. Fig. 5. Pitching moment load strain results.
changed theoretically, so the interference of thrust and drag loading on lift measurement is eliminated by electric method. As shown in the figure, the effective direction (parallel direction along the X axis) strain is 31.4 micro-strains and 31.3 micro-strains in No.21 and No.22 gauge position, respectively. The total resistance of the bridge arm formed by two strain gauges of the same type tends to zero after forming one bridge arm at the above two positions, which will have a good effect of elimi nating disturbance after post calibrating. For the interference of drag on pitching moment, as shown in Fig. 2, the thrust force can only make the force-moment decomposition mechanism of large triangle rotate as a whole. Theoretically, the torque measuring element does not deform. As shown in the simulation results of Fig. 3, only about 1.4 micro-strains are produced in the effective strain direction (along the Y-axis parallel di rection) of the pitch moment element patch position when the drag is loaded, which achieves a good effect of eliminating the interference. The main reason of the strain is that the moment reference center of the support is slightly shifted after the force is applied. As seen in Fig. 4, that lift measurement sensitive structure produce about 88 micro-strains in the effective strain direction of the strain gauge (parallel direction along the X-axis) when the support force measurement system is loaded with a full-scale load in the lift direction. The lift force has little influence on the thrust or drag measurement, and the effective direction of strain gauge application produces about 3 strain quantities of micro-strain magnitude. For the disturbance of lift to pitching moment, as that lift load position passes through the intersec tion point of the spatial concurrent force system, the lift force only cause the tension and compression deformation of the pitch moment measuring beam, and the strain at the patch position is in the same di rection and the absolute value is the same, so the pitch bridge theory does not cause unbalance, the output voltage does not increase, and the interference of the lift force on the pitch moment measuring is eliminated. As seen in Fig. 5, that pitch moment measurement sensitive structure produce about 93 microstrains in the effective strain direction (parallel direction along the Y axis) of the strain gauge when the support force measurement system is loaded with a full-scale load in the pitch moment
Fig. 4. Lift force load strain results.
structure of the measuring system. Lift measurement bridge adopts eight strain gauges combination bridge form. The two strain gauges positions of the same bridge arm are symmetrical about the Y axis centerline of the thrust or drag measuring element beam. When disturbance strain occurs, the two strain gauges resistances of the same bridge arm increase and decrease exactly one by one, and the absolute values are the same, which ensures that the resistance values of the four bridge arms are not 3
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Flow Measurement and Instrumentation 70 (2019) 101646
direction. The pitching moment has little influence on the thrust or drag measurement, and the effective direction of strain gauge application produces about 4.5 strain magnitudes. The interference of pitching moment to lift force is also eliminated by electrical method. Although the effective strain of each strain gauge is about 31 micro-strains, the strain directions of two strain gauges in the same bridge arm are opposite. Theoretically, the total resistance of each arm does not change, and there is no output voltage increment of the lift bridge. Table 1 shows the principal strain and the relative principal strain disturbance strain of three measurement directions under each compo nent individual full-scale loading. The strain is calculated as the strain of each bridge arm. The maximum disturbance is the disturbance of pitching moment loading to thrust or drag measurement, which is 3.9%. In impulse combustion wind tunnel test, the system will bear several times the design load of hypersonic airflow impact load [19], so the strength analysis must be carried out. The allowable stress is calculated according to the maximum possible transient load, and the allowable stress formula is shown in Equation (4).
σ � ½σ � ¼
σb
Fig. 6. Von-Mises stress cloud diagram.
(4)
nc ⋅na
Where σb is the yield stress of 00Ni18Co8Mo5TiAl, which is 1750 MPa nc and na represent the impact factor and safety factor of the wind tunnel, respectively. According to the test requirements, if the impact factor is 3 and the safety factor is 2, the design safety factor n ¼ nc � na ¼ 6. The maximum design stress σ should not be greater than the allowable stress [σ]. The allowable stress [σ] is 291.7 MPa. All components of the model are loaded synthetically, including the thrust/drag force is 2000 N, the lift force is 3000 N, and the pitching moment is 1000Nm. The calculated Von-Mises stress cloud diagram is shown in Fig. 6. It can be seen from the figure that the maximum stress concentration point appears on the pitching moment measuring beam structure under comprehensive loading. The stress value is 194 MPa, which is less than the allowable stress 291.7 MPa. The system meets the test requirements in strength. Because the effective force measurement time of the impulse com bustion wind tunnel used by the force measurement system is very short, the average force measurement time is only 250 ms. In order to ensure the accuracy of the force measurement results, at least 6 cycles of data shall be obtained during the test period [10,20]. According to the test requirements, the first natural frequency of the system shall reach 30 Hz. Therefore, modal analysis is carried out on the simplified structure of the model and the assembly of the support force measuring system, the model material is aviation hard aluminum 7075, and the total mass is about 200 Kg. Fig. 7 shows the first six modal modes of the force measuring system under the constraint condition at the bottom of the support. Table 2 shows the calculation results of the first six modal natural frequencies. The results show that the first order natural fre quency of the force measuring system reaches 37 Hz. The first vibration of the thrust drag, lift and pitching moment measured by the dyna mometer system occurs at the fourth order 204 Hz, the sixth order 298 Hz and the third order 78 Hz, respectively, which meets the test requirements.
Fig. 7. The first six modal modes. Table 2 First six natural frequencies and main vibration directions.
4. Conclusion In summary, we proposed a new type of hypersonic wind tunnel
Modal order
Frequency (Hz)
Main Vibration Direction
1 2 3 4 5 6
37.6 38.7 78.8 204.6 285.6 298.6
Lateral direction Yaw direction Pitching direction Thrust/Drag direction Alignment direction Lift direction
Table 1 Main strain and interference strain loaded by each component separately.
Thrust or drag Fx ¼ 2000 N Lift Fy ¼ 3000 N Pitching moment Mz ¼ 1000Nm
Thrust or drag strain
Interference with thrust or drag
Lift strain
Interference with lift
Pitching moment strain
Interference with pitching moment
�114με �3.1με �4.5με
\ �2.7% �3.9%
＜0.5με �88με ＜1.5με
0.6%＜ \ 1.7%＜
�1.4με �0 �93με
�1.5% �0% \
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Flow Measurement and Instrumentation 70 (2019) 101646
org/10.1016/j.flowmeasinst.2019.101646.
support force measuring system, simulated and analyzed the strain sensitivity, the interference between components and the modal attri butes of each measurement component of the support force measuring system. The results show that all the parameters meet the requirements of hypersonic wind tunnel load test. The design also provides a new idea for wind tunnel load test. The main advantages of the design are as follows:
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1. The support force measuring system changed the model-internal balance-support configuration in hypersonic wind tunnel test, the system structure composed of the model and the load-measuring support is formed, which does not need to design the internal bal ance separately and does not occupy the internal space of the test model, thus greatly facilitating the layout and design of the inte grated model with engines. The external position of the loadmeasuring structure also makes the installation and maintenance of the system more convenient. 2. Different from the traditional external balance, the special structure of the concurrent force system of the dynamometer system greatly reduces the disturbance of the additional moment caused by the thrust/drag and lift to the pitching moment. The finite element simulation results show that the additional moment produced by thrust force only makes the whole concurrent force system structure rotate, the pitching moment measuring element has almost no effective strain, the lift force only makes the tensile and compressive deformation of the pitching moment measuring beam, and the electrical decomposition of Wheatstone bridge makes the interfer ence between the measuring components very small. 3. The support force measuring system eliminates the bolt connection between the traditional balance and the support by integrating the traditional internal balance with the support. Therefore, the support force measuring system has greater stiffness, higher response fre quency and better linearity, and is more suitable for the instanta neous dynamic force measurement in hypersonic wind tunnel test. Acknowledge This work is financially supported by the National Natural Science Foundation of China (51775452 and 51805457) and the Fundamental Research Funds for Central Universities of China (2682019CX35, 2682017ZDPY09 and 2018GF02). Appendix A. Supplementary data Supplementary data to this article can be found online at https://doi.
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