Aeronautics and astronautics – Spacecraft – Attitude control
Reexamination Certificate
1999-10-19
2001-09-18
Jordan, Charles T. (Department: 3644)
Aeronautics and astronautics
Spacecraft
Attitude control
C267S164000, C248S557000
Reexamination Certificate
active
06290183
ABSTRACT:
BACKGROUND OF THE INVENTION
To reduce the dynamic loads transmitted to a spacecraft by its launch vehicle during launch, a vibration isolation system is desired which isolates and serves as the mounting system for the spacecraft to the launch vehicle. However, owners of spacecraft that cost tens to hundreds of millions of dollars demand a high strength, high fatigue-life connection between the spacecraft and the launch vehicle. This connection must provide a fail-safe connection; must be able to handle, without stress failure, the deflections due to the sum of the quasi-static acceleration loads of the spacecraft due to maneuvering and other vehicle loading events, and the dynamic loads of the isolation system; must be completely linear in all deflection regions (both tension and compression); and must be of minimal height and minimal weight. The isolation system must also not introduce collateral problems with the launch, such as low frequency modes, interaction with the launch vehicle control system, or reduction of payload-fairing clearance. The isolation system must be easily tunable for different combinations of launch vehicles and spacecraft, and readily employable in existing spacecraft because of the importance of flight heritage. Vibration isolation systems which provide isolation for all three translation axes are required where significant vibration loads exist in each of the axes. Particular to spacecraft, lateral dynamic loads during launch may often be as critical to the payload spacecraft as the vehicle longitudinal dynamic loads.
This invention is a device which simply and compactly provides passive vibration isolation in all three translation axes. When multiple devices are properly sized and configured together, a whole-spacecraft passive vibration isolation system with six degree-of-freedom passive vibration isolation is created. This system particularly provides a means for substantially reducing all the translational and rotational components of vibration transmitted to spacecraft from their launch vehicles during the launch process. By varying the size and spacing of the individual devices, the system can be easily tuned to suppress vibration loads at low or high frequencies. The system is also effective for substantial attenuation of shock loads. The benefits afforded the spacecraft and its components include reduced structural weight and cost, as well as increased life and reliability.
Historically, the connection between the spacecraft and the launch vehicle has been made with a very stiff structure. That type is generally considered to be a “hard mount” and is extremely efficient at transmitting all structure-borne forces from the launch vehicle to the spacecraft over a very wide frequency band. A need exists for isolating the payload of a launch vehicle from all structure-borne vibration loads; those due to launch, maneuvering, thrust termination and staging, as well as periodic thrust oscillations, pyrotechnic separation systems and aerodynamic loading.
Vibration isolation systems work by connecting the isolated structure (payload) to the base structure by means of a resilient mount or mounts. Damping is required in the resilient mounts to reduce the amplitude of response of the payload at the isolation frequency when the system is under external excitation at the isolation frequency. The resilient mounts must also allow sufficient relative motion between the vibrating base structure and the payload, which is referred to as the isolator stroke, or sometimes referred to as the “rattle space.”
Because the spacecraft is a major structural component of the launch vehicle/spacecraft dynamic system, variations in the isolation frequencies greatly effect the dynamics of the launch vehicle/spacecraft system. Any unpredicted changes in the dynamics can have an adverse effect on the control system of the launch vehicle and cause instability and thereby loss of the mission. Therefore, the stiffness properties of the isolation system must be accurately predicted and accounted for throughout the entire flight. The simplest and most effective way of achieving this predictable isolation system performance is by having a linear isolation system under all load cases, including launch vehicle acceleration loads, which typically range from −2 g's to +6 g's and higher. Resilient mounts commonly use a soft, non-linear material, such as an elastomeric, as their stiffness component. However, because of their non-linearity, elastomerics (rubbers, etc.) exhibit different stiffness under various loads, temperatures, and frequencies, resulting in complexity and unpredictability in performance, and therefore they cannot effectively be used as the stiffness component of a whole-spacecraft passive vibration isolation system. Also, under very high static loads, elastomerics creep (deflect as a function of time), and this cannot be tolerated. The use of elastomeric material as the stiffness component has been due to its heretofore advantage in tolerating strains up to 50%, which has allowed the elastomeric isolation mount to provide the necessary isolator stroke.
Three-axis whole-spacecraft vibration isolation design has eluded previous attempts. The disclosed invention, which is elegant and simple, satisfies all of these requirements.
SUMMARY OF THE INVENTION
The invention described within is a three-axis vibration isolation device effective for implementing a whole-spacecraft passive vibration isolation system providing substantial vibration load isolation in all three axes of any orthogonal coordinate system. The present invention incorporates the axial vibration isolation device of application Ser. No. 08/980,790 in multiplicity in a manner that produces a three-axis vibration isolation device that is suitable for three-axis whole-spacecraft vibration isolation.
A whole-spacecraft passive isolation system, to be effective and practical, must provide substantial isolation of the payload spacecraft from high frequency dynamic loads while simultaneously supporting the spacecraft under high G quasi-static acceleration loads with minimal movement of the payload spacecraft relative to the launch vehicle structure. A three-axis whole-spacecraft passive isolation system must satisfy these needs in all load directions. The device described does these things and retains the substantial improvements in dynamic load isolation, fatigue-life performance, and linear load-vs.deflection behavior achieved with the axial passive vibration isolation device referenced in application Ser. No. 08/980,790. The three-axis device is also very compactly configured so that it may be simply and straightforwardly employed in presently fabricated hardmount spacecraft support structures.
The three-axis isolation device disclosed achieves these qualities and improvements over other vibration isolation devices by capitalizing on the linear stiffness, and high strength of the axial vibration isolation device, as well as its substantially stiffness-independent damping characteristic. The flexure element of one axial device is connected to the flexure element of a second axial device in such a way as to create a predictable and linear lateral compliance within the new assembly. A stand-off post of high strength, linear-elastic material is moment connected to one of the flexure beams of the axial flexure element by way of the device mount located on that flexure beam. An axial device is thus connected at each end of the stand-off post thereby creating an assembly of two axial devices serially and moment connected by the stand-off post. The axial stiffness of the assembly is governed by the two axial devices connected serially, and is substantially determined by the simple transverse load, fixed-end bending stiffness of the flexure beams of the individual axial devices. The lateral stiffness of the assembly, however, is largely determined by the torsional stiffness of the flexure beams and also by the local moment bending stiffness of the flexure beams. Since the two axial devices are moment c
Johnson Conor D.
Pendleton Scott C.
Wilke Paul S.
Creighton Wray James
CSA Engineering, Inc.
Dinh Tien
Jordan Charles T.
Narasimhan Meera P.
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