AEROSPACE VEHICLE WEIGHT AND BALANCE ESTIMATION SYSTEM AND METHOD FIELD OF THE DISCLOSED TECHNIQUE The disclosed technique relates to aerospace vehicles in general, and to systems and methods for estimating weight and balance of aerospace vehicles, in particular. BACKGROUND OF THE DISCLOSED TECHNIQUE Knowledge of the weight of an aircraft is crucial for its operation, and safety. Knowledge of the weight allows the calculation of the maximum payload that can be transported a known distance and the amount of fuel required. Knowledge of the point of balance or center of gravity (CG) of an aircraft is also important. For example, if the longitudinal CG is located too forward, the aircraft will be nose heavy, if located too aft, tail heavy. A tail heavy aircraft that deviates from the recommended tolerances may become hazardously unstable exhibiting uncharacteristic spin and stall characteristics. Prior art methods for determining the weight of an aircraft include the use of aviation scales, and weighing sensors located in the landing gear, particularly in the main and nose wheels axles. Each of these prior art methods has its own disadvantages. The main function of the landing gear is not only to support the weight of the aircraft on the ground, but also to at least partly dissipate the tremendous amount of energy produced during the landing impact. Weighing sensors located in the landing gear, and especially the wheel axles are affected by changing conditions that include mechanical factors (e.g., damping pressure, tire pressure, vertical elasticity of the tire, material fatigue, etc.), environmental factors (e.g., temperature, humidity, corrosion, contaminants, etc.), variability in periodic maintenance and service, and the like. Hence, these regular changing conditions may hamper the reliability, accuracy, and efficacy of weight assessment. The relationship between weight and balance of an aircraft and its safety is recognized and documented. For example, an article entitled "Analysis of aircraft weight and balance related safety occurrences" published by the National Aerospace Laboratory (NLR) of the Netherlands, studies weight and balance related incidents (including accidents) of passenger as well as cargo aircraft. This study concludes that the accuracy and reliability of (then-known) prior art weight and balance systems are insufficient to impose their use as a primary means for determining aircraft weight and balance. U.S. Patent No.: US 8,235,326 B2, issued to Braincourt et al. and entitled "Aircraft Landing Gear Load Sensor" is directed at a fiber optic load sensing system and method for measuring load in an aircraft landing gear. The fiber optical load sensing system includes a plurality of Bragg Grating sensors written into a fiber optic cable, and an interrogator. The fiber optic cable is mounted, such that it is firmly clamped or bonded to the inside of an axle (right and/or left) of an aircraft landing gear. When the axle deflects under vertical and/or drag load, the optical fiber bends in sympathy. The interrogator determines the change in radius of the optical fiber caused by the bending. The change in geometry of the optical fiber is equated to the load that caused the deflection. A remote control and a recorder unit record the output of the fiber optic load sensing system. A plurality of load sensing measurements is taken corresponding to each wheel location such that the load apportionment and total load can be established for a wheel group of the aircraft. PCT International Publication Number WO 2015/088987 A1 to Moog Inc., entitled "Fiber Optic Sensing and Control System" is directed at a fiber optic sensing aeronautical flight control system for air and space vehicles. The fiber optic sensing and control system includes sensing optica! fibers, and an interrogator unit. The sensing optica! fibers each include multiple fiber optic sensing points that are integrated or coupled to primary and secondary flight control surfaces of the aircraft. The so-called primary flight control surfaces of the aircraft (airlerons, elevators, and rudder) are used to control aircraft movement in the pitch, yaw, and roll axes, whereas the so-called secondary flight control surfaces (inboard and outboard spoilers, inbound and outbound flaps, flaperons, and slats), are used to influence the lift or drag of the aircraft. The optical fibers are connected to the interrogator unit. The flight surfaces of the aircraft exhibit aeroelastic effects as wel! as structural loads, as gusts of wind and other forces are applied thereto. The sensing optical fibers sense and measure deformations and oscillations caused by the structural loads. The interrogator interrogates the sensing optical fibers. A flight control computer of the aircraft analyzes the measurements, the analyzed results of which are then fed back to an actuator that controls its corresponding flight control surface. SUMMARY OF THE PRESENT DISCLOSED TECHNIQUE It is an object of the disclosed technique to provide a novel method and system for estimating weight of an aerospace vehicle. In accordance with the disclosed technique, there is thus provided a weight estimation system for estimating weight of an aerospace vehicle while grounded. The weight estimation system includes a measurement subsystem and a processor. The measurement subsystem includes at least one sensor configured to measure a physical property (e.g., strain) in at least one of a fuselage, a wing, and an interface that interfaces at least one of the fuselage and the wing with an undercarriage of the aerospace vehicle, in at least one area exhibiting a measurable change in geometry (e.g., deformation) that is at least partly due to the weight. The measurement subsystem is configured to produce measured data indicative of the weight of the aerospace vehicle. The processor is configured for receiving at least part of the measured data and further configured to estimate the weight, by relating the measured data with predetermined physical-property-to-weight correspondence data associated with the aerospace vehicle. In accordance with the disclosed technique, there is further provided a method for estimating weight of an aerospace vehicle while grounded, the method includes the procedures of measuring a physical property in at least one of a fuselage, a wing, and an interface that interfaces at least one of the fuselage and the wing with an undercarriage of the aerospace vehicle, in at least one area exhibiting a measurable change in geometry that is at least partly due to the weight; producing measured data indicative of the weight of the aerospace vehicle, according to measured physical property; and estimating the weight by relating the measured data with predetermined physical-property-to-weight correspondence data associated with the aerospace vehicle. In accordance with the disclosed technique, there is further provided a sensor system for an aerospace vehicle. The sensor system includes at least one sensor coupled with at least one of a fuselage, a wing, and an interface that interfaces at least one of the fuselage and the wing with an undercarriage of the aerospace vehicle. The at least one sensor is configured to measure a physical property in at least one of the fuselage, the wing, and the interface, in at least one area exhibiting a measurable change in geometry that is at least partly due to the weight of the aerospace vehicle while grounded. The at least one sensor is further configured to produce measured data indicative of the weight of said aerospace vehicle. BRIEF DESCRIPTION OF THE DRAWINGS The disclosed technique will be understood and appreciated more fully from the following detailed description taken in conjunction with the drawings in which: Figure 1 is a schematic block diagram illustrating a weight and balance estimation system, constructed and operative in accordance with an embodiment of the disclosed technique; Figure 2 is a schematic illustration of an aerospace vehicle showing a plurality of measurement areas, selected in accordance with the embodiment of the disclosed technique; Figure 3A is a schematic illustration showing measurement areas and sensor placement about a main gear support structure of an exemplary aircraft, according to the principles of the disclosed technique; Figure 3B is a schematic illustration showing measurement areas and sensor placement on a fuselage section of an exemplary rotorcraft, according to the principles of the disclosed technique; Figure 3C is a schematic illustration showing an example of the applicability of the disclosed technique to another exemplary aircraft, according to the principles of the disclosed technique; Figure 3D is a schematic illustration of a bottom part, in greater detail, of the aircraft of Figure 3C, showing measurement areas and sensor placement on a fuselage section in proximity to a main landing gear of the aircraft, constructed and operative in accordance with the disclosed technique; Figure 3E is a schematic illustration showing an example interface, in greater detail, interfacing between fuselage section and an undercarriage of the exemplary aircraft shown in Figure 3D, constructed and operative in accordance with the disclosed technique; Figure 3F is a schematic illustration showing another example interface, in greater detail, interfacing between fuselage section and an undercarriage of the exemplary aircraft shown in Figure 3D, constructed and operative in accordance with the disclosed technique; Figure 3G is a schematic illustration of a bottom part, in greater detail, of aircraft of Figure 3C, showing measurement areas and sensor placement on a fuselage section in proximity to a nose landing gear of the aircraft, constructed and operative in accordance with the disclosed technique; Figure 4A is a schematic diagram illustrating an example method for compensating for temperature variation effects on strain measurements, constructed and operative in accordance with the disclosed technique; Figure 4B is a schematic diagram illustrating another example method for compensating for temperature variation effects on strain measurements, constructed and operative in accordance with the disclosed technique; Figure 4C is a schematic diagram illustrating a further example method for compensating for temperature variation effects on strain measurements, constructed and operative in accordance with the disclosed technique; Figure 5A is a schematic diagram showing a representative vehicle weight data acquirement technique in a vehicle-specific calibration method, constructed and operative in accordance with the embodiment of the disclosed technique; Figure 5B is a schematic diagram showing a calibration graph of the thermal strain as a function of temperature, under a constant weight of aerospace vehicle; Figure 5C is a schematic diagram showing a calibration graph of the mechanical strain as a function of weight of the aerospace vehicle being under constant temperature; Figure 5D is a schematic diagram showing exemplary vehicle-specific strain-to-weight correspondence data, constructed through the vehicle-specific calibration method, in accordance with the embodiment of the disclosed technique; Figure 6, which is a schematic illustration showing a method for estimating balance of the aerospace vehicle, constructed and operative in accordance with the embodiment of the disclosed technique; Figure 7A is a schematic block diagram of a preliminary physical-property-to-weight calibration method, constructed and operative in accordance with the embodiment of the disclosed technique; Figure 7B is a schematic block diagram of a method for estimating weight of an aerospace vehicle, constructed and operative in accordance with the embodiment of the disclosed technique; Figure 7C is a schematic block diagram of a method for estimating balance of the aerospace vehicle, constructed and operative in accordance with the embodiment of the disclosed technique; Figure 8A is a schematic illustration of a fiber-optic strain measurement subsystem, constructed and operative according to another embodiment of the disclosed technique; Figure 8B is a schematic illustration showing an example embedment of FBG optical fibers to a structural part of a fuselage section or wing section of the aerospace vehicle, which is subject to deformation; Figure 8C is a schematic illustration showing another example embedment of FBG optical fibers to another structural part of a fuselage section or wing section of the aerospace vehicle, which is subject to deformation; Figure 8D is schematic illustration showing a particular aspect in the operating principles of the strain measurement subsystem of Figure 8A; and Figure 9 is a schematic illustration showing an implementation of the weight and balance estimation system for determining a flight-ground status of the aerospace vehicle, DETAILED DESCRIPTION OF THE EMBODIMENTS The disclosed technique overcomes the disadvantages of the prior art by providing a weight estimation system and method for estimating the weight of an aerospace vehicle while grounded. The weight estimation system includes a measurement subsystem, and a processor. The measurement subsystem includes at least one sensor (e.g., a strain sensor, pressure sensor, etc.) configured to determine a physical property (e.g., strain, pressure) in at least one of a fuselage (e.g., a fuselage section), a wing (e.g., a wing section), and an interface that interfaces at least one of the fuselage and the wing with an undercarriage (landing gear) of the aerospace vehicle, in at least one area exhibiting a measurable change in geometry (e.g., deformation, displacement, configuration, surface shape) that is at least partly due to the weight of the aerospace vehicle. The measurement subsystem is configured to produce measured data indicative of the weight of the aerospace vehicle. The processor is configured and operative to receive at least part of the measured data and to estimate the weight, by relating the measured data with predetermined physical-property-to-weight (e.g., strain-to-weight) correspondence data associated with the aerospace vehicle. The method for estimating weight of the aerospace vehicle while grounded includes the steps of determining a physical property (e.g., strain, pressure) in at least one of a fuselage, a wing, and an interface that interfaces at least one of the fuselage and the wing with an undercarriage of the aerospace vehicle, in at least one area exhibiting a measurable change in geometry (e.g., deformation, displacement) that is at least partly due to the weight, producing measured data indicative of the weight of the aerospace vehicle, according to the measured physical property (e.g., strain), and estimating the weight by relating the measured data with predetermined physical-property-to-weight (e.g., strain-to-weight) correspondence (e.g., calibration) data associated with the aerospace vehicle. In other words, for each aerospace vehicle whose weight is to be determined by the weight and balance estimation system, there corresponds respective physical-property-to-weight (e.g., strain-to-weight) correspondence data that was predetermined in a preliminary calibration phase. The disclosed technique further discloses, in the context of a system, a sensor system (e.g., the measurement subsystem) for an aerospace vehicle (for acquiring measurements for estimating weight of the aerospace vehicle). The sensor system (i.e., which is onboard the aerospace vehicle) includes at least one sensor coupled with at least one of a fuselage, a wing, and an interface that interfaces at least one of the fuselage and the wing with an undercarriage of the aerospace vehicle. The at least one sensor is configured to measure a physical property (such as strain, pressure) in at least one of the fuselage, the wing, and the interface, in at least one area exhibiting a measurable change in geometry (such as deformation) that is at least partly due (typically predominately) to the weight of the aerospace vehicle while grounded. The at least one sensor is configured to produce measured data indicative of the weight of the aerospace vehicle. The produced measured data is transmitted to a processor (i.e., which can be onboard, off-board the aerospace vehicle, or embodied partly onboard and partly off-board). The processor is configured to receive the measured data, and to estimate the weight of the aerospace vehicle by relating the received measured data with predetermined physical-property-to-weight (e.g., strain-to-weight) correspondence (e.g., calibration) data associated with the aerospace vehicle. The term "aerospace vehicle" used herein refers to a vehicle that is designed, intended or capable of travel (e.g., flying, traversing a distance) in at least one of: (1) a fluid (e.g., typically gaseous) atmosphere (e.g., Earth's atmosphere), (2) a vacuum (e.g., space, atmosphere-less environment). Examples of aerospace vehicles that fly in (e.g., Earth's) atmosphere include variable-wing and fixed-wing aircraft (e.g., monoplanes, biplanes, gliders, commercial, military and private aircraft, etc.), rotorcraft (e.g., helicopters, autogyros, cyclogyros, compound rotorcraft, etc.), aerodynes (e.g., vertical and/or short take-off and landing (V/STOL) such as the Harrier, V-22 Osprey, etc.), unmanned aerial vehicles (UAVs), airships (e.g., dirigibles, blimps, Zeppelins, etc.), ornithopter (i.e., a wing flapping aircraft), and the like. Examples of aerospace vehicles that travel in space include spacecraft such as landers (crafts) (e.g., lunar lander, Mars lander, etc.), artificial satellites, and the like. Examples of aerospace vehicles that travel both in the atmosphere and in space include spaceplanes (e.g., Space Shuttle, X-37 Orbital Test Vehicle (OTV)), launch vehicles (e.g., rockets such as expendable launch vehicles, reusable rocket vehicles, vertical takeoff horizontal landing (VTHL) vehicles, etc.), and the like. The term "grounded" used herein refers to situations where the aerospace vehicle is on a surface or ground (e.g., Earth, moon, Mars, etc.) being stationary (e.g., at rest or parked), moving (e.g., taxying, being towed, being transported), or at least partly suspended from the surface or ground (e.g., by a crane) in which the surface or ground is at least one of an area of land, at least one floating entity on a body of fluid (e.g., an aircraft carrier at sea), and at least partly carried by another entity such as another aerospace vehicle that is on a surface or in flight (e.g., aircraft transported within another aircraft, for example a helicopter transported within a military or cargo aircraft, a space shuttle carried on top of a transport aircraft, etc.). The term "physical property" used herein refers to a property of a physical entity that is measurable. Examples of physical properties are strain, pressure, electrical conductivity, electrical resistance, electric potential, magnetic flux, magnetic field, capacitance, inductance, electromagnetic wave related properties (e.g., optica! related properties such as radiance, luminance, refractive index, etc.), and the like. The disclosed technique in general, involves measurement of a physical property in particular areas, and in at least one structural member of an aerospace vehicle (i.e., excluding the undercarriage itself) that exhibits or experiences a measureable change in its geometry (e.g., preferably where maximal) at least partly (typically predominately) due the weight of the aerospace vehicle while grounded. For the purpose of elucidating the disclosed technique, and without loss of generality, some parts will be described in terms of a particular selected physical property (e.g., strain), although the principles of the disclosed technique likewise apply (i.e., with appropriate adaptations, where applicable) to the measurement of other physical properties (e.g., pressure). The phrase "change in geometry" refers to a change in a shape, configuration, or form of an object (e.g., a structural member of the aerospace vehicle). Example changes in geometry include deformation or strain (e.g., as a result of a force (e.g., weight) being applied to the object, pressure, a relative change in configuration or displacement of an object (e.g., a piston, displacement measurement sensor, etc.), and the like. The change in geometry (e.g., deformation due to the weight of the aerospace vehicle) is assessed with respect to a reference (i.e., a reference geometric state, such as a reference deformation state). Hence, a current geometric state (e.g., current deformation state) is assessed with respect to a reference geometric state (e.g., a reference deformation state). A reference geometric state (e.g., of a structural member) may be chosen arbitrarily, for example, corresponding to when the aerospace vehicle is at its manufacturer's empty weight (MEW)), corresponding to when aerospace vehicle is at a known calibration weight, and the like. Other reference geometric states that may be chosen include specific sensor outputs that measure (directly or indirectly) the geometric state (e.g., a particular strain sensor output value, a camera acquiring images of a geometric state of an aerospace vehicle structural member, and a processor determining the change in geometry via image processing techniques). When the weight of aerospace vehicle changes (e.g., due to weight-loading) there is a corresponding change in geometry such as weight-induced deformation (e.g., of at least one structural member such as a fuselage section, wing section and an interface that interfaces between at least one of the fuselage section and wing section with an undercarriage). Deformation is generally defined as a transformation of an object from one configuration (e.g., reference configuration) to another configuration. Reference is now made to Figure 1, which is a schematic block diagram illustrating a weight and balance estimation system, generally referenced 100, constructed and operative in accordance with an embodiment of the disclosed technique. Weight and balance estimation system 100 includes a measurement subsystem 102, a processor 104, a memory 106, a communication subsystem 108, and a user interface 110. Measurement subsystem 102 includes at least one sensor, typically a plurality thereof, designated sensors 1121; 1122,...,112N (where subscript N denotes a positive integer index). Processor 104 is coupled with measurement subsystem 102, memory 106, communication subsystem 108, and user interface 110. Alternatively, or additionally, communication subsystem 108 is coupled with measurement subsystem 102 (denoted by dotted line 116). Weight and balance estimation system 100 is configured and operative to be communicatively coupled with an external weight and balance calibration system 114 (i.e., not part weight and balance estimation system 100) for the purpose of calibration. Communication subsystem 108 is configured and operative to communicatively couple with external weight and balance calibration system 114 for the duration of the calibration procedure, as will be elaborated. Alternatively, or additionally, processor 104 is configured and operative to communicatively couple with externa! weight and balance calibration system 114 for the duration of the calibration procedure (as represented by dotted line 118). Reference is now further made to Figure 2, which is a schematic illustration of an aircraft showing a plurality of physical property measurement areas, selected in accordance with the embodiment of the disclosed technique. Without loss of generality, Figure 2 illustrates an example of an aerospace vehicle 130 having the general form of a commercial aircraft (and particularly, in this example, a Boeing® mode! 747-8) that includes a fuselage 132, main wings 1341, 1342, and landing gears 1361, 1362, 1363, 1364, and 1365. Landing gears 1361, 1362, and 1363 are coupled with different areas of fuselage 132, while landing gears 1364 and 1365 are each respectively coupled with areas of main wings 134-i and 1342. Further without loss of generality, the physical property selected to explicate the principles of the disclosed technique is strain and the physical property measurement areas will be described and interchangeably referred to as "strain measurement areas" or for brevity "measurement areas". Particularly, further shown is a plurality of strain measurement areas 1381, 1382 (located on fuselage 132), 1383, and 1384 (located respectively on main wings 1341 and 1342). Generally, a physical property measurement area is generally defined as an area on the aerospace vehicle where the physical property is measured. Specifically and for example, a strain measurement area is generally defined as an area on the aerospace vehicle where strain is measured. In particular, the physical property (e.g., strain) that is measured according to the principles of the disclosed technique is primarily caused by minute changes in geometry (e.g., deformations) of the airframe (i.e., excluding the landing gear) in areas that bear the most weight while the aerospace vehicle is grounded. The amount of change in the geometry is associated with and is relative to the amount of weight of the aerospace vehicle. For comparison, in an extreme case, the change of geometry is relative to a situation where aerospace vehicle 130 is in a weight-less environment (i.e., does not experience its weight (or acceleration)). Generally, the areas on the fuselage and/or wings of an aerospace vehicle typically bearing the most weight are those that couple (directly or indirectly) to, or in the vicinity of, the landing gears. Hence, measurement areas (e.g., strain measurement areas) are typically selected on at least one fuselage section and/or wing section and/or an interface that interfaces at least one of the fuselage section and wing section with the undercarriage of the aerospace vehicle that is adjacent, or in close proximity to the landing gears. An interface may be embodied, for example in the form of a coupler (e.g., a coupling member such as a shaft, mechanical fasteners, piston, adhesive, etc.), a thin structure (e.g., a sheet-like structure that interfaces between at least one of a fuselage section and wing sections with an undercarriage section), and the like. Among other requirements (e.g., support of aerodynamic pressure distribution), the structure of aerospace vehicle 130 is required to support (i.e., resist and transmit) two different types of applied loads: ground loads (i.e., exhibited during ground movement (e.g., taxying, towing)) and aerodynamic loads (i.e., exhibited during flight). When aerospace vehicle 130 is grounded (e.g., parked) its undercarriages or landing gears as well as its structure function to distribute and bear the weight required to be supported (including the weight of the flight crew, passengers, cargo, fuel, supplies, etc.). For the sake of clarity, the main structures of a typical aerospace vehicle are the fuselage, wings, and landing gear (also interchangeably referred herein as "undercarriage"). Various parts of aerospace vehicle 130, for example, interfaces such as structural members, and couplings that interface and/or couple the landing gear with the fuselage and/or main wing sections exhibit greater weight bearing loads than other more distant parts (in relation to the undercarriage), such as the tail section or the vertical stabilizer structure. The structures and couplings ("structural elements") coupling the landing gear with the fuselage and/or main wing sections are viable areas for coupling sensors (e.g., strain sensors) for measuring changes in the geometry (e.g., deformation) of those areas, predominately due to the weight of the aerospace vehicle. A typical coupling configuration (interface) between landing gears and the body (i.e., fuselage and wings) of the aerospace vehicle includes a trunnion (part of landing gear) that is configured to engage and couple (e.g., via trunnion pins - interface members that function as couplers) with landing gear support beams that are fixed to the fuselage or wing sections of the aerospace vehicle. The interface members may also function as dampers (e.g., springs, pistons). Reference is now made to Figure 3A, which is a schematic illustration showing measurement areas and sensor placement about a main gear support structure of an exemplary aircraft, according to the principles of the disclosed technique. Figure 3A illustrates a main gear support configuration of an exemplary aircraft in which the undercarriage (landing gear) 140 is partly supported by a fuselage section 142 and partly by a wing section 144 (wing box). A gear wing beam 146 pivotally couples undercarriage 140 with fuselage section 142 (at one end) as well as with wing section 144 (at another end). A side brace 148 couples undercarriage 140 with fuselage section 142, and a lower drag brace 150 couples undercarriage 140 with an upper drag brace 152, which in turn is coupled with wing section 144. Figure 3A further shows a plurality of strain measurement areas 1385, 1386, 1387, as well as a plurality of strain sensors 1123, 1124, 1125 1126, 1127, 1128. Particularly, strain measurement area 1385 is located on fuselage section 142, such to include the area where a part of gear wing beam 146 and a part of side brace 148 are coupled with and are at least partially supported by fuselage section 142. Strain measurement area 1386 is located on wing section 144, such to include the area where the other part of gear wing beam 146 is coupled with and at least partially supported by wing section 144. Strain measurement area 1387 is located on wing section 144, such to include the area where upper drag brace 152 is coupled with and is at least partially supported by wing section 144. Strain sensors 1123 and 1124 are coupled with fuselage section 142, and are located in strain measurement area 1385 in proximity to the coupling of gear wing beam 146 to fuselage section 142. Strain sensors 1125 and 1126 are coupled with wing section 144, and located at strain measurement area 1387 in proximity to the coupling of upper drag brace 152 to wing section 114. Strain sensors 1127 and 1128 are coupled with wing section 144, and are respectively located on strain measurement areas 1387 and 1388 in proximity to the coupling of gear wing beam 146 to wing section 144. Strain measurement areas 1385, 1386, 1386 are located where a significant amount of the weight of undercarriage 140 (i.e., and therefore at least part of aerospace vehicle 130 being distributed among the landing gears) is supported by fuselage section 142 and wing section 144 (i.e., when aerospace vehicle 130 is grounded and undercarriage 140 is in an extended (non-retracted) position). Consequently, fuselage section 142 and wing section 144 typically exhibit measurable deformation that is predominately due to at least the partial weight of aerospace vehicle 130 being supported. Likewise, any interface (such as a mechanical coupler, fastener, thin sheet-like structure, pivot pin, etc) that interfaces between an undercarriage (landing gear) and fuselage section 142 and/or wing section 144 may experience strain that is at least partly due to the weight of aerospace vehicle 130 when grounded. As such, the interface itself may be considered as having at least one strain measurement area (not shown). It is noted that the interface itself may be considered part of the fuselage and/or part of the wing and/or part of the undercarriage. It is further noted that according to one implementation the interface itself is constructed, configured, and operative to replace an existing interface (having no sensing capabilities) such as a pivot pin, mechanical coupler, etc. (and concurrently function the same as the interface it was intended to replace). According to another implementation the interface itself is constructed, configured, and operative to be retrofitted to an existing interface. Strain sensors 1123, 1124, 1125 1126, 1127, 1128 are configured and operative to sense and measure this measurable deformation as strain exhibited by their associated strain measurement areas, corresponding to fuselage section 142 and wing section 144. From the individual strain measurements produced by the strain sensors, measurement subsystem 102 is configured and operative to determine the strain (i.e., at least one strain value) of at least one of fuselage section 142 and wing section 144 (typically both) and to produce strain data indicative the weight of aerospace vehicle 130. Measurement subsystem 102 also determines strain in fuselage and/or wing sections associated with or in proximity to the remaining landing gears of aerospace vehicle 130 (e.g., nose landing gear, left/right side main landing gears, and the like). Alternatively, the sensors are embodied instead as pressure sensors (not shown) that are positioned at measurement areas in a manner that enables measurement of at least part of the pressure generated between at least one of fuselage section 142, wing section 144, and an interface that interfaces at least one of fuselage section 142 and wing section 144 with undercarriage 140, at least partly due to the weight of aerospace vehicle 130. Further alternatively, the sensors are embodied as cameras (not shown), which are configured and operative to acquire measured data in the form of images (not shown) of at least one of fuselage section 142, wing section, and an interface that interfaces at least one of fuselage section 142 and wing section 144 with undercarriage 140 in at least two weight states (e.g., a reference weight state and a current weight state of aerospace vehicle 130) of aerospace vehicle 130. Processor 104 is configured and operative to receive these images from the cameras acquired at the two weight states (e.g., the reference weight state can be a reference calibration weight) and to determine a change in geometry (e.g., deformation) in a structural member of aerospace vehicle 130 corresponding to the two weight states. As will be described below in greater detail, based on the measured data from the sensors, processor 104 is configured to estimate the current weight of aerospace vehicle 130. There is an assortment of coupling methods for coupling the sensors with fuselage section 142 and wing section 144. Examples include the use of mechanical fasteners, as well as adhesives such as thermosetting plastics (e.g., epoxy cements), cyanoacrylate cement, ceramic cements, polyester epoxies, and the like. The use of a particular coupling material and method has to ensure compatibility such that the coupling material allows proper transmission of the measured physical property such as strain or pressure from the measured object (specimen or sample) and the sensor (e.g., strain sensor, pressure sensor, respectively). The coupling material's physical and chemical properties are also taken into account, as it may possess a different coefficient of expansion with respect to the specimen and/or sensor. For instance, if the coupling material is an adhesive, its curing process is also taken into account, as the adhesive may experience temperature induced expansion, contraction caused by cooling, exhibit residual internal stresses, as well as a phenomenon of post-cure shrinkage, all of which may influence strain measurements. The principles of the disclosed technique apply to various types of aerospace vehicles. To further demonstrate the multi-platform applicability, reference is now further made to Figure 3B, which is a schematic illustration showing measurement areas and sensor placement on a fuselage section of an exemplary rotorcraft, according to the principles of the disclosed technique. Figure 3B illustrates the case where aerospace vehicle 130 is a rotorcraft, such as helicopter 160. Figure 3B (top) illustrates an exemplary rotorcraft (helicopter) 160 that includes a fuselage 162 and an undercarriage 164 (embodied in the form of landing skids, and fore and aft struts 1641, 1642). Undercarriage 164, which is coupled with a bottom part of fuselage 162, supports the weight of helicopter 160 while grounded. The bottom illustration of Figure 3B, which is a partial enlargement of the top illustration, shows a plurality of strain measurement areas 1387, 1388, 1389, 13810 that are located on a bottom portion of fuselage 162, and which correspond to the locations where fuselage 162 exhibits the largest amount of strain when helicopter 160 is grounded. Particularly, (transverse) strain measurement areas 1388 and 13810 respectively encompass the entire span of fore and aft struts 1641; 1642, as the latter support the weight load of helicopter 160. Longitudinal strain measurement areas 1387 and 1389 extend at least partially along the two bottom longitudinal sides of fuselage 162 section. It is to be understood that different types, makes, and models of aerospace vehicles in general, and helicopter 160 in particular, would have different strain measurement areas, which depend, at least in part, on structure, configuration, and the specific nature of the coupling between the undercarriage and fuselage and/or wing sections whose areas exhibit (typically the most substantial) measurable deformation due to the weight of aerospace vehicle 130. In accordance with a particular embodiment of the disclosed technique, and without loss of generality, measurement subsystem 102 employs electro-optical strain measurement techniques (e.g., Brillouin scattering sensing techniques, Rayleigh scattering sensing techniques, interferometric sensing techniques, Bragg grating sensing techniques, etc.). According to one example, measurement subsystem 102 includes at least one sensor embodied in the form of a fiber Bragg grating (FBG) sensor that employs an optical fiber 166 possessing at least one periodic variation in the refractive index of its fiber core. Optical fiber 166 is coupled with an onboard or off-board interrogator (represented by arrow 168). A description disclosing greater detail of such an embodiment will be provided hereinbelow in conjunction with Figures 8A-8D. Generally, at least part of the optical fiber 166 of the FBG sensor is coupled with (e.g., incorporated with, embedded into) a section of fuselage 162 along a path traversing at least part of strain measurement areas, as exemplariiy shown in Figure 3B. Measurement subsystem 102 employing the FBG sensor is configured and operative to determine strain in at least one strain measurement area in general, and typically a plurality of discrete strain measurement points or intervals along the optical fiber path in particular. In an alternative configuration, there is a plurality of FBG sensors each configured to determine strain in different or at least partially along overlapping paths (not shown). The coupling of optical fiber 166 to fuselage 162 may involve a diversity of methods, such as the use of adhesives (acrylics), and thermosetting polymers (e.g., polyimide), etc. Reference is now made to Figures 3C, 3D, 3E, 3F, and 3G. Figure 3C is a schematic illustration showing an example of the applicability of the disclosed technique to another exemplary aircraft, according to the principles of the disclosed technique. Figure 3D is a schematic illustration of a bottom part, in greater detail, of the aircraft of Figure 3C, showing measurement areas and sensor placement on a fuselage section in proximity to a main landing gear of the aircraft, constructed and operative in accordance with the disclosed technique. Figure 3E is a schematic illustration showing an example interface, in greater detail, interfacing between fuselage section and an undercarriage of the exemplary aircraft shown in Figure 3D, constructed and operative in accordance with the disclosed technique. Figure 3F is a schematic illustration showing another example interface, in greater detail, interfacing between fuselage section and an undercarriage of the exemplary aircraft shown in Figure 3D, constructed and operative in accordance with the disclosed technique. Figure 3G is a schematic illustration, in greater detail, of a bottom part of aircraft of Figure 3C, showing measurement areas and sensor placement on a fuselage section in proximity to a nose landing gear of the aircraft, constructed and operative in accordance with the disclosed technique. Figures 3C, 3D, and 3G illustrate an example of aerospace vehicle 130 being a UAV (drone) 180 that includes a fuselage 182, wings 184, main landing gear (MLG) 186, and a front (nose) landing gear (NLG) 188. MLG 186 and NLG 188 are of the non-retractable type. The principles of the disclosed technique likewise apply to landing gears that are of the retractable type (e.g. as shown in Figure 3A). Figure 3D shows portion 190 (Figure 3C) of UAV 180 in greater detail. In particular, Figure 3D illustrates a section of fuselage 182 and a section of MLG 186 decoupled (separated) from each other, thereby highlighting strain measurement area 13811 being located on a bottom part of fuselage 182 that is configured to couple with MLG 186. Further shown is a rigid coupling base 194, a plurality of MLG coupling members 196i, 1962, 1963, and 1964 (not shown), a plurality of fuselage coupling members 1981, 1982, 1983, and 1984, and an FBG strain sensor 200 (i.e., partly embodied as an optical fiber). Rigid coupling base 194 is coupled (braces) with MLG 186, as well as with MLG coupling members 1961, 1962, 1963, and 1964 (i.e., via two longitudinal struts that at least partly form the rigid coupling base structure). Fuselage coupling members 1981, 1982, 1983, and 1984 are coupled with a bottom portion of fuselage 182, and are further configured and operative to couple (mate, pair) with corresponding MLG coupling members 1961, 1962, 1963, and 1964 at respective complementary positions. FBG strain sensor 200 is coupled with a bottom section of fuselage 182 in strain measurement area 13811 that exhibits deformation predominately due to the weight of UAV 180, The example traversal circuit of FBG strain sensor 200 shown in Figure 3D is along a path which enables multiple strain sensing points or intervals. Other paths within strain measurement area 13811 are viable (not shown). FBG strain sensor 200 is coupled with other components of measurement subsystem 102 (represented by arrows 202 and 204) as will be described in greater detail hereinbelow in conjunction with Figures 8A-8D. Figure 3E partially shows circled portion 206 (Figure 3D) of UAV 180 in greater detail. In particular, Figure 3E shows an example interface 210 (embodied in the form of a coupler, e.g., a coupling shaft), that is configured and operative to interface between fuselage section 182 and undercarriage 186 of the exemplary aircraft 180 shown in Figure 3D, and further configured and operative to include (e.g., incorporate, involve) and/or be associated with at least one sensor 212 (i.e., corresponding to at least one sensor 1121,...,112N that is part of measurement subsystem 102 - Figure 1) for measuring a physical property (e.g., strain) exhibited at least partly as a result of the weight of UAV 180. The sensor associated with interface 210 is typically further coupled with a communication subsystem 214, which in turn is configured and operative to receive (e.g., strain) measurements from the (at least one) sensor and to provide these measurements to processor 104 (e.g., via communication subsystem 108). The particular example of interface 210 shown in Figure 3E relies on a relative change in displacement (d) of parts thereof, so as to produce a signal that is at least partly dependent of the weight of UAV 180. Generally, sensor 212 measures a physical property in an area exhibiting a measurable change in geometry (e.g., displacement, change in configuration) at least partly due to the weight of UAV 180. A change in the geometry (e.g., of the interface) causes a corresponding change to the physical property being measured. Sensor 212 can be embodied, for example, in the form of a displacement sensor, pressure sensor, capacitance sensor (e.g., in which a change in the geometry (e.g., displacement) causes a corresponding change in capacitance), an inductance sensor, a strain sensor, an electrical resistance sensor, an electrical conductance sensor, a magnetic flux sensor, a magnetic field sensor, a general electromagnetic wave related sensor (e.g., optical based sensors), and the like. Figure 3F shows another example interface 216 (embodied in the form of a 2-D manifold or sheet-like structure), that is configured and operative to interface between an area of fuselage section 182 and an area of undercarriage 186 of UAV 180, and further configured and operative to include (e.g., incorporate, involve) and/or be associated with at least one sensor 218 (i.e., corresponding to at least one of sensors 1121,..., 112N - Figure 1) for measuring a physical property, such as pressure or strain exhibited at least as a result of the weight of UAV 180. The particular example of interface 216 relies on strain being applied to the interface at both sides thereof (i.e., at one side that interfaces fuselage section 182, as well as the other side that interfaces the undercarriage 186). The example shown in figure 3F shows that sensor 218 associated with interface 216 is selected to be an FBG sensor, however, other types of sensors likewise apply (e.g., resistive strain sensors, capacitive strain sensors, semiconductor strain sensors, pressure sensors, and the like). Figure 3G shows portion 192 (Figure 3C) of UAV 180 in greater detail. Figure 3G illustrates a section of fuselage 182 and a section of NLG 188. A rigid coupling base 220 couples NLG 188 with fuselage 182. NLG 188 may typically include an oleo strut 222. A strain measurement area 13812 extends such to include the area where rigid coupling base 220 couples with fuselage 182. A strain sensor 224 (e.g., an FBG type, partly shown as an optica! fiber) is coupled with a bottom section of fuselage 182 in strain measurement area 13812 that exhibits deformation predominately due to the weight of UAV 180, An example path of optical fiber 224 within strain measurement area 13812 is shown in Figure 3G. Strain sensor 224 is coupled with measurement subsystem 102 (as represented by arrow 226). Alternatively, other paths of optical fiber 224 are viable (e.g., a rectangular path, elliptical path, etc.). Further alternatively, multiple strain sensors (of same or different type) may be employed (not shown). In alternative implementations, and without loss of generality, strain sensor 200 is instead a resistive stain sensor, capacitive strain sensor, inductive strain sensor, semiconductor strain sensor, acoustical strain sensor, mechanical strain sensor, piezoelectric strain sensor, etc. Further alternatively, at least two strain sensors are employed and are of different type (not shown). Measurement subsystem 102 (Figure 1) is configured and operative to determine (i.e., may also be referred interchangeably herein as measure, assess, estimate) strain in strain measurement areas 1381, 1382, 1383, and 1384 (i.e., by measuring a physical property correlated with strain, such as electrical resistance outputted from a resistive-type strain gauge). In particular, sensors 1121, 1122,...,112N of measurement subsystem 102 are distributed among, and installed at, measurement areas 1381, 1382 1383, and 1384. Installation of the sensors can be achieved by a variety of methods. For example, according to a first sensor installation method, sensors 1121, 1122,...,112N are integrated (at least partially or entirely) into the structure of fuselage 132 and sections of main wings 1341 and 1342 (Figure 2) in measurement areas 1381, 1382 1383, 1384 (Figure 2) during the manufacturing of aerospace vehicle 130. According to a second sensor installation method, sensors 1121, 1122,...,112N are retrofitted to an existing aerospace vehicle 130 in measurement areas 1381, 1382 1383, and 1384. According to a third sensor installation method, only part of sensors 1121, 1122,...., 112N are installed in the manufacturing process of aerospace vehicle 130 in measurement areas 1381, 1382, 1383, and 1384, while at least another part of sensors 1121, 1122,..., 112N is retrofitted in measurement areas 1381, 1382 1383, and 1384 sometime following the manufacturing of aerospace vehicle 130. Sensors 1121, 1122,...,112N of measurement subsystem 102 are configured to measure a physical property correlated with a physical property such as strain in areas exhibiting measurable elastic deformation (i.e., strain measurement areas 1381, 1382 1383, and 1384) that is at least (typically predominately) due to the weight of aerospace vehicle 130 while it is grounded. Measurement subsystem 102 is configured and operative to produce strain data indicative of the strain and also that is indicative of the weight of aerospace vehicle 130 while grounded. There are various ways in which the disclosed technique acquires strain measurements. In accordance with one strain measurement technique, sensors 1121, 1122,...,112N are embodied as strain sensors, also interchangeably referred herein as "gauges" that are configured and operative to exhibit, produce or output a measurable physical property, such as electrical conductance (or conversely resistance - in resistive strain gauges), capacitance (in capacitive strain gauges), inductance (in inductive strain gauges), etc, as a function of applied strain or changes in the sensor's geometry. Without loss of generality, for elucidating the principles of the disclosed technique, we will now arbitrarily select to describe particular aspects of the disclosed technique, by way of example, such that the sensor is a strain type sensor and the measured physical property of the strain sensor is electrical resistance. A strain sensitivity factor S (gauge factor) of each of strain sensors 1121, 1122,...,112N (in the case they are embodied as the resistive strain gauge type) is known and generally given by: where R0 is the electrical resistance of the strain gauge when unstrained, AR is the change in the electrical resistance of the strain gauge subjected to applied strain, and ε is the strain. There are, however, other effects, apart from plain mechanical strain that influence the strain measurement. Example effects that may influence the strain measurement include the thermal expansion (or contraction) of the measurement object whose strain is measured, the temperature-dependence of the strain gauge, the temperature dependence of the electrical conductors (connecting wires), etc. Hence equation (1) represents a simplistic case where temperature effects are not taken into account. Weight and balance estimation system 100 is constructed, operative and intended for use with aerospace vehicles located in different environments varying in temperature. The disclosed technique takes into account such temperature-dependent effects on the measured strain. In general, the effect of temperature on the resistance of the strain gauge may be given by: where a is the temperature coefficient of the strain gauge, AT is the change in temperature and ΔR/R0 is a unit change in resistance from an initial reference resistance R0 caused by a change in temperature. To measure strain of an object such as a fuselage section or a wing section of the aerospace vehicle, strain sensors are typically coupled with (e.g., embedded into or onto) the object that is measured, so that at least part of the strain is transmitted from the strained object to the strain sensor. In such a case, equation (2) may be modified to take into account temperature-induced strain effects of the object: where α0 is the thermal expansion coefficient of the object. The newly added term in equation (3), namely (α0_α) . ΔT, vanishes if the thermal expansion coefficient of the strain gauge material is made to match that of object, which is one method of minimizing temperature effects on strain measurements, as will be elaborated hereinbelow. The disclosed technique takes into account (i.e., corrects, compensates) the effect of temperature on the physical property being measured in general, and specifically as described by way of example on strain measurements in a variety of methods. The disclosed technique may employ self-temperature compensated strain sensors, as well as those that are non-self-temperature compensated. Generally, in non-self-temperature compensated strain sensors, a change in temperature of the strain sensor will produce a corresponding change in its thermal output (whereby the temperature-induced measurement change is independent of the stress-induced mechanical strain). Self-temperature compensated strain sensors, in contrast, are devised to exhibit minimal thermal output. In the example of resistive-type strain gauges that are non-self-temperature compensated, a temperature change of AT will produce a corresponding change in resistance AR. To at least partially counter the thermal output effect, which may be considered a significant source for strain measurement errors, according to one method of compensating for temperature variation effects on strain measurements, self-temperature compensated strain gauges may be employed, which may be constructed from, for example, sundry constantan alloys that are selected so as to match the material of the aerospace vehicle whose strain is to be measured. According to another compensation method for temperature variation effects on strain measurements, the measurement subsystem includes a plurality of temperature sensors each thermally coupled with a respective sensor, for measuring its temperature. To further elucidate the particulars of this method, reference is now further made to Figure 4A, which is a schematic diagram illustrating an example method for compensating for temperature variation effects on strain measurements, constructed and operative in accordance with the disclosed technique. Figure 4A illustrates a particular embodiment, denoted 1021, of the measurement subsystem shown in Figure 1, constructed and operative to compensate (correct) for temperature variation effects on strain measurements, in addition to strain gauge sensors 1121, 1122,..,112N, measurement subsystem 1021 further includes a plurality of N separate temperature sensors 2301, 2302,...,230N each thermally coupled (T) to a respective strain sensor (according to index number). In other words, temperature sensor 2301 is thermally coupled with strain sensor 1121, temperature sensor 2302 is thermally coupled with strain sensor 1122, and so forth. Each temperature sensor 2301, 2302,...,230N is configured and operative to sense the temperature of its respective strain sensor 1121, 1122,...,112N and to output respective temperature data. Measurement subsystem 1021 uses the temperature readings (data) outputted by each of the temperature sensors to compensate for the strain measurements outputted from sensors 1121, 1122,...,112N, according to a temperature compensation model (e.g., based on equation (2)). According to a further compensation method for temperature variation effects on strain measurements, the measurement subsystem includes a plurality of temperature sensors that are each integrated into a respective strain sensor. To further elucidate the particulars of this compensation method reference is now further made to Figure 4B, which is a schematic diagram illustrating another example method for compensating for temperature variation effects on strain measurements, constructed and operative in accordance with the disclosed technique. Figure 4B illustrates another particular embodiment, denoted 1022, of the measurement subsystem shown in Figure 1, whereby each strain sensor 112'1, 112'2,...,112'N includes and incorporates a respective temperature sensor 2321, 2322,...,232N. Hence, strain sensor 112'i includes temperature sensor 2321, strain sensor 112'2 includes temperature sensor 2322, and so forth. The temperature sensors are integrated into the construction of the strain sensors, such to measure their temperatures, so as to facilitate measurement subsystem 1021 in compensating for temperature variation effects on strain measurements (e.g., according to a temperature compensation model, based for example on equation (2)). According to another compensation method for temperature variation effects on strain measurements, the measurement subsystem includes complementary (secondary) sensors to the (primary) measurement sensors that function to at least partially reduce (e.g., minimize, cancel) the temperature variation effects of the primary measurement sensors. To further elucidate the particulars of this compensation method reference is now further made to Figure 4C, which is a schematic diagram illustrating a further example method for compensating for temperature variation effects on strain measurements, constructed and operative in accordance with the disclosed technique, Figure 4C illustrates a further particular embodiment, denoted 1023, of the measurement subsystem shown in Figure 1. Specifically, measurement subsystem 1023 includes strain measurements sensors 112"1, 112"2,...,112"N (also denoted herein interchangeably as "primary", and which are identical respectively to strain measurement sensors 112-j, 1122,...,112IM of Figure 1), as well as complementary (also denoted herein interchangeably as "dummy" or secondary) sensors 2341, 2342,...,234N to the primary strain measurement sensors. Hence, for each primary strain measurement sensor 112"1, 112"2,...,112"N there is a respective secondary or complementary sensor associated or paired therewith. In other words, strain measurement sensor 112"1 is paired with complementary sensor 2341, strain measurement sensor 112"2 is paired with complementary sensor 2342, and so forth to 112"N and 234N. Each complementary sensor is thermally coupled with a sample not subject to strain that is comprised from a material identical to the object whose strain is measured by its corresponding primary strain measurement sensor. Given that each pair of primary and secondary sensors are affected in the same manner to variations in temperature, measurement subsystem 1023 is configured to use the strain-induced measurement outputted by a primary sensor and the unstrained measurement outputted by the dummy sensor, both being at the same temperature, so as to derive a temperature compensated strain measurement. Alternatively, processor 104 is configured and operative to receive the strain-induced measurement outputted by the primary sensor and the unstrained measurement outputted by the dummy sensor, so as to derive a temperature compensated strain measurement. For example, in resistive-type strain gauges, given that the strain factor is known for both the primary and secondary strain sensors, processor 104 and/or strain measurement subsystem 102 solves equation (2) for ε, by measuring or determining the change in the electrical resistance, ΔR, at specifically measured temperatures. According to a further temperature compensation method for temperature variation effects on strain measurements, weight estimation system 100 is configured and operative to filter out the effect of slowly changing temperature variations (i.e., when temperature changes relatively slowly with respect to the entire duration of the weight estimation procedure). Specifically, processor 104 filters out the effect of slowly changing temperature variations, by appiying filtering techniques, such as by employing a low-pass filter, and the like. In general, and in a similar manner to temperature, environmental factors (other than temperature, such as humidity, wind speed, solar radiation, etc.) that may influence measurements may be compensated for by various techniques. An example for one such technique employs the aforementioned primary and secondary (dummy) sensor approach, where the primary sensor is configured to sense a physical property (e.g., strain), while the secondary sensor is independent of the physical property and configured to sense the environmental factor that is to be compensated for. For such purposes, weight and balance estimation system 100 may further include additional environmental measurement sensors, such as at least one humidity sensor (not shown), at least one anemometer (not shown), at least one radiation detector (not shown), and the like. The environmental measurements acquired by these sensors are provided (e.g., transmitted) to processor 104 for processing, and for countering the environmental effects on the physical property being measured. Weight and balance estimation system 100 includes multiple modes of operation. According to one mode of operation, measurement subsystem 102 acquires measurements of a physical property (e.g., strain, pressure), produces corresponding measured data, and transmits the measured data to processor 104, which in turn determines a measured property value, such as strain e (e.g., from equation (3)). According to another mode of operation, measurement subsystem 102 acquires measurements, produces corresponding measured data, at least partially processes the measured data, and transmits the at least partially processed measured data to processor 104, which in turn determines a measured physical property (e.g., strain ε). According to a further mode of operation, measurement subsystem 102 acquires measurements, produces corresponding measured data, and processes the measured data to determine the physicai property (e.g., strain E). For each mode of operation, there exists a sub-mode for temperature compensation. According to one sub-mode, measurement subsystem 102 acquires temperature measurements (from at least one temperature sensor), produces corresponding temperature data, transmits the temperature data to processor 104, which in turn reduces temperature variation effects on physical property (e.g., strain) measurements. According to another sub-mode, measurement subsystem 102 acquires temperature measurements, produces corresponding temperature data, at least partially processes the temperature data into partially processed temperature effect compensation data, and transmits the at least partially processed temperature effect compensation data to processor 104, which in turn reduces (e.g., cancels) temperature variation effects on the physical property (e.g., strain) measurements. According to a further sub-mode, measurement subsystem 102 reduces temperature variation effects on acquired measurements (without substantive intervention of processor 104). The disclosed technique includes a vehicle-specific calibration method that relates strain measurement data with different reference calibration weight measurements acquired at a reference temperature-controlled environment of the vehicle. This vehicle-specific calibration method involves several phases. Without loss of generality, the physical property selected to elucidate the calibration method is strain. To elucidate the vehicle-specific calibration method, reference is now further made to Figures 5A, 5B, 5C, and 5D. Figure 5A is a schematic diagram showing a representative vehicle weight data acquirement technique in a vehicle-specific calibration method, constructed and operative in accordance with the embodiment of the disclosed technique. Figure 5B is a schematic diagram showing a calibration graph of the thermal strain as a function of temperature, under a constant weight of aerospace vehicle. Figure 5C is a schematic diagram showing a calibration graph of the mechanical strain as a function of weight of the aerospace vehicle being under constant temperature. Figure 5D is a schematic diagram showing exemplary vehicle-specific strain-to-weight correspondence data, constructed through the vehicle-specific calibration method, in accordance with the embodiment of the disclosed technique. The vehicle-specific calibration method described in greater detail below is facilitated, at least in part, by processor 104 (Figure 1) processing the acquired calibration data. Alternatively, according to another approach of the disclosed technique, the processing of the calibration data is achieved (wholly or in part) by at least one processor different (not shown) from processor 104, such as a remote computer, an on-site computer, via distributed computing methods, and the like. Figure 5A illustrates aerospace vehicle 130 situated in a temperature-controlled environment 240 that maintains a temperature T (that can be varied). Figure 5A further shows one representative example of a weighing technique known in the art, employed by the disclosed technique for calibration purposes. According to a first phase in the calibration method, aerospace vehicle 130 is configured (e.g., arranged, placed, mounted in a level position) for weighting on an external (separate) weight and balance calibration system 114 (e.g., entailing aviation scales) that may include one or typically a plurality of weight measurement elements: a weight determination processor 1141, and a plurality of platform scales 1142, 1143, and 1144. External weight and balance calibration system 114 may also include a level subsystem (not shown) to ensure that aerospace vehicle 130 is maintained at a known attitude (e.g., a zero level position) throughout the calibration procedure. As shown in Figure 5A, the wheels of aerospace vehicle 130 are placed correspondingly on platform scales 1142, 1143, and 1144, which in turn are coupled with weight determination processor 1141 Platform scales 1142, 1143, and 1144 sense the weight that is applied thereto and generate corresponding signals, which in turn are communicated to weight determination processor 1141 that interprets the signals to produce a corresponding output (termed interchangeably as "calibration weight data" or "calibration weight measurement") indicative of the weight of aerospace vehicle 130. It is noted that without loss of generality, the disclosed technique may utilize other types of weighing systems and techniques for calibration, such as jack weighing techniques (not shown), landing gear weight measurement techniques, etc. External weight and balance calibration system 114 provides (e.g., transmits, communicates) the calibration weight data to processor 104 directly 118 (Figure 1) or indirectly (e.g., via communication system 108). Processor 104 is configured and operative to receive the calibration weight data from external weight and balance calibration system 114 and to store this data in memory 106. Once the weight of aerospace vehicle 130 is known from the aforementioned first phase, the vehicle-specific calibration method further involves a second phase of determining or isolating the temperature contribution or effect on the measured physical property. Particularly, in the present example, the contribution or effect of thermal strain on strain measurements that are acquired by sensors 1121,...,112N (i.e., as opposed to mechanical strain predominately due to the weight of aerospace vehicle 130). In general, the total strain experienced by an object involves a superposition (algebraic sum) of the mechanical strain ΕM (i.e., describing shape changes, deformation, or relative displacement of particles in an object, resulting from mechanical stresses) and the thermal strain εT (i.e., strain due to thermal effects, such as thermal expansion and contraction, thermal output (e.g., electrical resistivity, thermal expansion differential between signal conductor and test conductor, existent for example, in certain resistive-type strain gauges), etc.). In particular, processor 104 is configured and operative to determine the thermal strain under constant weight of aerospace vehicle 130. Initially, processor 104 receives the calibration weight data from external weight and balance calibration system 114 (and may also monitor that the weight of aerospace vehicle 130 remains substantially constant during the calibration procedure). In addition, processor 104 is configured and operative to receive respective outputs from sensors 1121, 1122,...,112N (resistance measurements) and to calculate ΔR/R0 (given a known initial reference resistance of R0) at different temperatures T of temperature-controlled environment 240, such that ΔT= T-Tref (where Tref. denotes a known reference temperature). To this end, the temperature T of temperature-controlled environment 240 is capable of being altered (e.g., with respect to the reference temperature), the value T of which is provided to processor 104 (e.g., via communication system 108). For a plurality of values of T (or ΔT for that matter) and their respective values of ΔR/R0, processor 104 solves equation (2) for 8 as well as for constants S and α, the latter two of which generally depend on the strain gauge material(s). Once constants S and α are determined, and assuming a constant weight of aerospace vehicle 130, the contribution of the thermal strain to the total strain measurement is known. Figure 5B illustrates a representative calibration graph 242 of thermal strain εT as a function of temperature (T) assuming aerospace vehicle 130 maintains a constant weight (W). A third phase of the calibration procedure involves determining the contribution of the measured physical property (excluding the thermal effect). Particularly, in the present example the procedure involves determining the contribution of mechanical strain (excluding the contribution of thermal strain) as a function of the weight of aerospace vehicle 130 being under constant temperature. In this phase, the temperature of aerospace vehicle 130 within temperature-controlled environment 240 is kept constant (e.g., Tref), while the weight of aerospace vehicle 130 is methodically varied (e.g., by progressively adding calibration weights of known value to the base weight (e.g., manufacturer's empty weight (MEW)) of aerospace vehicle 130). Measurement subsystem 102 determines the mechanical strain as a function of the weight of aerospace vehicle 130 measured by external weight and balance calibration system 114. Figure 5C shows a calibration graph 244 of the mechanical strain as a function of weight of the aerospace vehicle being under constant temperature. In a fourth phase of the calibration procedure, given the vehicle-specific behavior of the thermal effect (e.g., thermal strain) as a function of temperature under constant weight (determined in the second phase), as well as the physical property (e.g., mechanical strain) excluding the thermal strain as a function of weight under constant temperature (determined in the third phase), processor 104 is configured and operative to determine total physical-property-to-weight (e.g., strain-to-weight) correspondence (e.g., calibration) data associated with (e.g., specific to) aerospace vehicle 130. With reference to Figure 5D, processor 104 constructs a database (e.g., lookup table) of vehicle-specific physical-property-to-weight (strain-to-weight) correspondence data 246 that relates the physical property (e.g., total determined strain) as a function of calibration-determined weight. In other words, for each total strain determined value εi, there exists a corresponding weight value Wi (where i represents an index between 1 and m, and m is a positive integer representing the number of strain-to-weight pair values in the database). In general, for each physical property determined value, there exists a corresponding weight value, thereby forming physical-property-to-weight correspondence data, which associates the physical property value with a corresponding weight value. Memory 106 (Figure 1) is configured and operative to store the strain-to-weight correspondence data 246. The vehicle-specific calibration method is applicable and is performed to each specific vehicle or type of vehicle (e.g., of the same make and model) whose weight estimation is required during operation of weight estimation system 100. Following calibration, during operation of weight and balance estimation system 100, for a particular aerospace vehicle (e.g., 130) whose weight is required to be ascertained while grounded, and whose specific physical-property-to-weight (e.g., strain-to-weight) correspondence data is known or predetermined via the calibration method, there are two main modes of operation (i.e., for weight estimation). The first main mode of operation does not require the determination of the temperature of strain sensors 1121,...,112N as well as the object whose strain is to be measured (e.g., fuselage section, wing section). According to this approach, measurement subsystem 102 employs temperature compensated sensors and/or methods as described for example hereinabove, in conjunction with Figures 4A-4C. Measurement subsystem 102 is configured to determine strain in at least one of a fuselage section (or plurality thereof), a wing section (or plurality thereof) of aerospace vehicle 130, and an interface that interfaces between at least one fuselage section and wing section with an undercarriage of aerospace vehicle 130, in at least one area exhibiting measureable change in geometry (e.g., deformation) that is at least partly (or typically predominately) due to its weight, and to correspondingly produce (sampled) strain data. Processor 104 receives the sampled strain data from measurement subsystem 102, and estimates the weight of aerospace vehicle 130 by relating the determined (sampled) strain data with the strain-to-weight correspondence (e.g., calibration) data associated with aerospace vehicle 130. This association may typically be specific (i.e., a particular aerospace vehicle), partly-specific (i.e., parameters related to the type, make, model, series, age, flight history, etc. of the aerospace vehicle), and the like. For example, for a determined (sampled) strain value E' there exists a corresponding and closest matching database strain values (i.e., one of strain values ε1,... ,εm in strain-to-weight correspondence database 246). In general, the strain-to-weight correspondence data may typically be indicative of the behavior of a monotonic increasing function, whereby increasing strain values measured are indicative correspondingly of increasing weight. Processor 104 relates the determined (measured) physical property (strain value ε') with one of a corresponding and closest matching database physical property (strain value ε) (e.g., via extrapolation), which in turn is associated with a respective weight value (determined in the calibration method). Thus, weight and balance estimation system 100 provides a weight estimation value for aerospace vehicle 130 by relating currently measured physical property measurement data (e.g., strain data) with a database weight value of aerospace vehicle 130, according to a predetermined relationship of physical-property-to-weight (e.g., strain-to-weight) correspondence data specific to aerospace vehicle 130. According to a second main mode for weight estimation, measurement subsystem 102 employs non-self-temperature compensated strain sensors. According to this approach, weight and balance estimation system 100, in general, and measurement subsystem 102 in particular, include a plurality of temperature sensors (e.g., temperature sensors 1801,...,180N, or 1841,...,184N - Figures 4A-C) configured and operative to measure the temperature of their respective strain sensors, as well as the temperature of the object whose strain is measured (i.e., structural elements in the fuselage section(s) and/or wing section(s) and/or at least one interface that interfaces at least one of a fuselage section and wing section with the undercarriage of aerospace vehicle 130). Measurement subsystem 102 is configured and operative to calculate the change in temperature AT between the measured temperature (i.e., outputted by the temperature sensors) and a known reference temperature Tref. (e.g., T=const., Figure 5C), to further measure and calculate ΔR/R0, so to determine the strain ε' (sampled strain data). Similarly, to the first mode of operation for weight determination, processor 104 receives the sampled strain data from measurement subsystem 102 and estimates the weight of aerospace vehicle 130 by relating the measured sampled strain data e' with the strain-to-weight correspondence data (i.e., a closest matching database strain value E (i.e., one of strain values ε1,..., εm), which in turn corresponds with an associated weight value (i.e., one of weight values W1,...,Wm)), associated with aerospace vehicle 130. According to another aspect of the disclosed technique, weight and balance estimation system 100 is configured and operative to estimate balance of aerospace vehicle 130 while grounded. Balance is an important factor that affects the safety, operability, and efficiency of aerospace vehicle 130 during flight. Generally, an improperly balanced aircraft could result in reduced or impaired controllability of the aircraft during flight, which could possibly lead to an accident or damage. Additionally, an improperly balanced (e.g., nose-heavy, tail-heavy) aircraft would entail expending more energy in the form of engine power and consequently fuel, in order to maintain the aircraft in level flight. Knowing an aircraft's center of gravity (abbreviated herein "CG" or interchangeably "CoG", defined as the average location of the weight of the aircraft (where it is balanced)) in relation to its center of lift (abbreviated herein "CoL", defined as the location where the sum total of all lift is generated or considered to be concentrated) is crucial in determining its stability and controllability, The disclosed technique employs the determined estimated weight of aerospace vehicle 130 to derive an estimation of its point of balance or determined CG as will be elaborated in greater detail in the following description. Reference is now made to Figure 6, which is a schematic illustration showing a method for estimating balance of the aerospace vehicle, constructed and operative in accordance with the embodiment of the disclosed technique. As shown in Figure 6, aerospace vehicle 130 maintains a constant orientation with respect to ground 250. The procedure of estimating the balance involves determining if aerospace vehicle 130 is level (i.e., horizontal, or "level flight attitude"). An external level (e.g., electronic spirit level - not shown) coupled with (e.g., mechanically, optically) aerospace vehicle 130 is configured and operative to determine the attitude or level value of aerospace vehicle 130, produce corresponding attitude data. This attitude data is provided to processor 104 (e.g., level transmits attitude data to processor 104 via communication subsystem 108 (Figure 1)). Alternatively, the inertia! navigation system (INS) of aerospace vehicle 130 determines the attitude or level value. The attitude data (e.g., level value) is provided to processor 104 (e.g., manually via user interface 110 (Figure 1), automatically via communication subsystem 108, etc.). Further alternatively, weight and balance estimation system 100 includes a level (not shown), coupled with processor 104, configured and operative to determine the attitude or level value of aerospace vehicle 130. Balance estimation involves determining moments of force (torques) with respect to a reference datum (also denoted interchangeably herein as "datum"). The datum is an imaginary vertical reference plane from which horizontal distance measurements are made or computed for the purpose of balance estimation. Figure 6 shows a typical (e.g., Boeing©) reference datum 252, located in front of the nose of aerospace vehicle 130, as well as another reference datum 2522. Without loss of generality, we select to use reference datum 2522 for describing the principles of balance estimation according to the disclosed technique. Generally, for each aircraft, the position of different components is identified with respect to the datum. The datum is typically indicated in the Aircraft Specifications, Spacecraft Specification, etc. The positions (coordinates) of sensors 1121,...,112N are also known or measured with respect to datum 2522. Strain sensors 1121,...,112N are positioned at various respective strain measurement areas 1381, 1382 1383, and 1384 (Figure 2). With reference to Figures 2 and 6A, measurement subsystem 102 is arranged and configured as a plurality of distinct and distanced apart sensor clusters 1021, 1022, 1023, and 1024 that are distributed at respective strain measurement areas 1381, 1382 1383, and 1384. Hence, the sensors are grouped in sensor clusters that are distributed among the (e.g., strain) measurement areas. In particular, sensor cluster 102, includes sensors 1121,..., 112i (where i≤1