To find the eigenvalues of a matrix, we solve the characteristic equation: \[ \det(A - \lambda I) = 0 \] where \( A \) is the matrix, \( \lambda \) is the eigenvalue, and \( I \) is the identity matrix.
Step 1: Construct the matrix \( A - \lambda I \)
Given the matrix \( A = \begin{bmatrix} 1 & 2 \\ 0 & 3 \end{bmatrix} \), we subtract \( \lambda I \): \[ A - \lambda I = \begin{bmatrix} 1 - \lambda & 2 \\ 0 & 3 - \lambda \end{bmatrix} \] Step 2: Find the determinant of \( A - \lambda I \)
\[ \det(A - \lambda I) = (1 - \lambda)(3 - \lambda) - (0)(2) = (1 - \lambda)(3 - \lambda) \] Step 3: Solve the characteristic equation
\[ (1 - \lambda)(3 - \lambda) = 0 \Rightarrow \lambda = 1 \text{ or } \lambda = 3 \] Conclusion:
The eigenvalues of the matrix are 1 and 3.
Let \[ A = \begin{pmatrix} 1 & 0 & 1 \\ 0 & k & 0 \\ 3 & 0 & -1 \end{pmatrix}. \] If the eigenvalues of \( A \) are -2, 1, and 2, then the value of \( k \) is _.
(Answer in integer)
A single-stage axial compressor, with a 50 % degree of reaction, runs at a mean blade speed of 250 m/s. The overall pressure ratio developed is 1.3. Inlet pressure and temperature are 1 bar and 300 K, respectively. Axial velocity is 200 m/s. Specific heat at constant pressure, \( C_p = 1005 \, {J/kg/K} \) and specific heat ratio, \( \gamma = 1.4 \). The rotor blade angle at the outlet is __________ degrees (rounded off to two decimal places).
An ideal ramjet with an optimally expanded exhaust is travelling at Mach 3. The ambient temperature and pressure are 260 K and 60 kPa, respectively. The inlet air mass flow rate is 50 kg/s. Exit temperature of the exhaust gases is 700 K. Fuel mass flow rate is negligible compared to air mass flow rate. Gas constant is \( R = 287 \, {J/kg/K} \), and specific heat ratio is \( \gamma = 1.4 \). The thrust generated by the engine is __________ kN (rounded off to one decimal place).
A monopropellant liquid rocket engine has 800 injectors of diameter 4 mm each, and with a discharge coefficient of 0.65. The liquid propellant of density 1000 kg/m³ flows through the injectors. There is a pressure difference of 10 bar across the injectors. The specific impulse of the rocket is 1500 m/s. The thrust generated by the rocket is __________ kN (rounded off to one decimal place).
Air at temperature 300 K is compressed isentropically from a pressure of 1 bar to 10 bar in a compressor. Eighty percent of the compressed air is supplied to a combustor. In the combustor, 0.88 MJ of heat is added per kg of air. The specific heat at constant pressure is \( C_p = 1005 \, {J/kg/K} \) and the specific heat ratio is \( \gamma = 1.4 \). The temperature of the air leaving the combustor is _______ K (rounded off to one decimal place).
An ideal turbofan with a bypass ratio of 5 has core mass flow rate, \( \dot{m}_a,c = 100 \, {kg/s} \). The core and the fan exhausts are separate and optimally expanded. The core exhaust speed is 600 m/s and the fan exhaust speed is 120 m/s. If the fuel mass flow rate is negligible in comparison to \( \dot{m}_a,c \), the static specific thrust (\( \frac{T}{\dot{m}_a,c} \)) developed by the engine is _________ Ns/kg (rounded off to the nearest integer).