BETA AIR, LLC

An electrically powered aircraft is configured estimate a true angle-of-attack (AoA) for an aircraft. A controller may receive one or more measurements associated with an aircraft, such as vane angle values, inertial measurement unit (IMU) data, and airspeed data. Discrepancies between vane angle values measured by different AoA sensors and an aircraft's true AoA may include fuselage effects, such as flow interference around the nose at different combinations of true AoA and angle-of-sideslip (AoS) causing different local AoA at the different vanes. The controller may generate and utilize a side slip estimate to determine if the difference in vane angle values is due to side slip or a malfunctioning AoA sensor. In some cases, if the difference in the vane angle values is determined to be due to a malfunctioning AoA sensor, then the controller may perform one or more actions (e.g., deactivate a stall barrier protection)

In an aspect of the present disclosure a pilot display system for an aircraft includes a sensor configured to measure an aircraft position and generate a position datum based on the aircraft position, a display incorporated in the aircraft, and a computing device communicatively connected to the sensor and the display, the computing device configured to receive a flight path including a transition point, determine, from the position datum, the aircraft position relative to the transition point, and command the display to display a visual representation of the aircraft position relative to the transition point.

An aircraft monitoring system for an electric aircraft is disclosed. The monitoring system may include at least a sensor configured to generate a failure datum. The failure datum includes a datum regarding a condition of an electric motors. An electronic checklist may include a crew alerting system (CAS). The CAS may be in electronic communication a sensor, wherein the CAS is comprised of at least a computing device. The CAS may generate a plurality of remedy data as a function of the failure datum. The CAS then may display the plurality of remedy data using a pilot display. A pilot may be prompted to apply the remedy datum using a pilot display as a function of the plurality of remedy data. The CAS provides an indication of the condition of the electric motor using a pilot display as a function of the application of the plurality of remedy data.

A system for establishing a primary function display for an electrical vertical takeoff and landing aircraft. The system further includes a plurality of sensors that detects at least a metric and generates at least a datum based on the at least a metric. Specifically, state of charge is at least generated based on the performance metric of an energy source. The system further includes a display to show the at least a datum. The system further includes a controller that receives the at least a datum and generates a visual to the pilot.

Systems and techniques for state control of a flight component of an electric aircraft that includes a high voltage connection to a power supply of the electric aircraft and a communication connection to a communication system of the electric aircraft. The flight component is configured with a hardware circuit designed and configured to receive and/or read a signal on the communication system, compare an identifier of the signal against a stored identifier of the flight component, and in response to the identifier matching the stored identifier, cause the flight component to change from a first state to a second state.

A system for monitoring sensor reliability in an electric aircraft is provided. The system includes a computing device communicatively connected to a first sensor and an electric aircraft. The first sensor is mechanically connected to the electric aircraft and is configured to detect a first flight datum of the electric aircraft. The computing device is configured to receive the first flight datum from the first sensor, compare the first flight datum to at least a corroboratory datum, and tag the first sensor as a function of the comparison of the first flight datum and the at least a corroboratory datum. A method for monitoring sensor reliability in an electric aircraft is also provided.

In an aspect of the present disclosure a pilot display system for an aircraft includes a sensor configured to measure an aircraft position and generate a position datum based on the aircraft position, a display incorporated in the aircraft, and a computing device communicatively connected to the sensor and the display, the computing device configured to receive a flight path including a transition point, determine, from the position datum, the aircraft position relative to the transition point, and command the display to display a visual representation of the aircraft position relative to the transition point.

A system for determining remaining useful energy in an electric aircraft, the system including a computing device where the computing device is configured to measure a internal state datum of a battery as a function of at least a sensor, receive the internal state datum from the at least a sensor, generate a useful energy remaining datum as a function of the internal state datum and a battery model, and display the useful energy remaining datum to a user.

A system for flight control configured for use in an electric aircraft includes an inertial measurement unit (IMU) and configured to detect an aircraft angle and an aircraft angle rate. The system includes a flight controller including an outer loop controller configured to receive the input datum from the sensor, receive the aircraft angle from the IMU, and generate a rate setpoint as a function of the input datum. The system includes an inner loop controller configured to receive the aircraft angle rate, receive the rate setpoint from the outer loop controller, and generate a moment datum as a function of the rate setpoint. The system includes a mixer configured to receive the moment datum, perform a torque allocation as a function of the moment datum, and generate a motor command datum as a function of the torque allocation.

Systems and methods relate to electric propulsor fault detection. An exemplary system includes at least a first inverter configured to accept a direct current and produce an alternating current, a first propulsor, a first motor operatively connected with the first propulsor and powered by the alternating current, and at least a noise monitoring circuit electrically connected with the direct current and configured to detect electromagnetic noise and disengage the at least an inverter as a function of the electromagnetic noise.