Considered is the method of ambiguity resolution for satellite-based GPS phase measurements, which takes into account the specific character of the problem of spacecraft relative navigation. This method provides for obtaining a fixed and a floating solution. Distinction of the proposed fixed solution consists in its layout into two stages and in more efficient arrangement of the integer search of phase measurement periods. At the first stage of the fixed solution the most probable value of the phase period integer is retrieved. At that, the exhaustive search of the period integer is not performed. High speed of this stage implementation is provided by the fact that the domain, in which searching is executed, is being continuously narrowed as far as new "candidates" for a more probable value are found. The second stage serves for estimating a posterior probability of a determined value of the period integer. It is essential that the estimation of a posterior probability is formed step-by step in the course of exhausting searching for possible values of the period integer. Moreover, the value of an estimate of posterior probability monotonely decreases while new integer values are being registered. Owing to such a method it becomes possible to abandon an uncertain fixed solution without complete calculation of its posterior probability and exhaustive search of all possible phase period values.

The method for a vehicle model parameters identification under active control providing for its motion periodic nature has been proposed. It has been shown that using the vehicle motion features some additional equations can be obtained that connect the parameters being identified. These equations can be effectively used when estimating the model parameters under external disturbances condition. The application of this approach to the parameters identification of a survey vessel model has been presented.

Using spherical trigonometry method, the equation was obtained which describes the law of generation of the aircraft commanded glide slope α_{���(=ILU)} (materialized as a green ray of indicating light) relative to the rocking aircraft-carrier deck. A block diagram of algorithm for generation of the stated angle is worked out. This angle provides stabilization of the prescribed glide slope α_�(=GS) relative to the deck averaged position with account for the aircraft-carrier static angles of trim and roll. Then the commanded glide slope α_� and hook-to-eye distance compensation angle of the aircraft are smoothly followed according to cosine law. The use of the obtained law of generating the current angle α_��� commanded to stabilization drive increases the glide slope stabilization accuracy, decreases deflection of the aircraft hook to aircraft-carrier deck touchdown point from the designed one and, therefore, increases safety of the aircraft landing upon the aircraft-carrier.

The paper presents the basic concept and results of INS servo system diagnostic system development. The diagnostic system consists of functional and test parts which are realized in all of modes of operation. The functional diagnostic in gyroscopic stabilization mode is based on instrument control of dynamic misalignment and informational control of INS invariant behavior - cosine angle of ESG angular momentums. The test control corresponds control of shift velocity tryout in servo system command channel. The concept of two-level informational control is justified. Such informational control must show instrumental failures and informational abnormalities in INS operation.

The results of development of a pendulum pulse-rebalance accelerometer unit (AU) calibration procedure is presented. For AU used on mobile objects with a smaller acceleration range, development of a procedure for calibration in the AU operating range of angles is relevant.
For calibration the AU is mounted on a single-axis turn-table which provides the possibility for the AU to be fixed in two opposite positions at an angle of 45^o relative to the axis of rotation.
The distinctive features of the proposed calibration technique are as follows:

• calibration is performed in the operating range of acceleration for exclude the errors due to the AU model inadequacy. That is typical for the pulse-rebalance accelerometer when calibrated in the range of ±g,
• the parameters are evaluated jointly with the AU installation errors using the statistical estimation methods to smooth over the measurement errors,
• an iterative procedure is used to exclude the model inadequacy in the case of large installation errors.

AU interchangeability assurance method for gimbal inertial navigation system is considered. The experimental results presented provide support for the efficiency of the proposed calibration technique.