Correct Answer : Homogeneous
Explanation : The working substance and chemical products should be homogeneous. When we say a substance is homogeneous, we intend to say that its composition and character is uniform throughout the substance.
Correct Answer : 0%
Explanation : An ideal rocket doesn’t allow any heat transfer to take place across its walls. The flow, in this case, would be adiabatic. So, 0% of heat transfer takes place across the walls of an ideal rocket.
Correct Answer : 1.8 kg/m3
Explaination : Ideal rockets have working substances that follow perfect gas law. Hence using P=ρRT, we can determine the density of the combustion product.R = Ru / Mmol = 8314 / 24 = 346.42 J kg-1K-1ρ = P/RT = 1 x 106 / (346.42 x 1600)= 1.8 kg/m3.
Correct Answer : Propellant flow is steady but constant
Explanation : In an ideal rocket, propellant flow is considered to be steady and constant. In such a vehicle, all kinds of frictional effects and boundary layers are neglected along with the possible discontinuities in flow such as shock waves. The expansion of the working fluid is both steady as well as uniform.
Correct Answer : Higher transient time
Explanation : An ideal chemical rocket engine has high pressure, high temperature and low flow velocity. The transient time is very low to ensure that the flow and associated processes are steady and remains constant at all times.
Correct Answer : 0.85
Explanation : The propellant mass fraction is the ratio of propellant mass to the initial mass of the vehicle.Minert = Mstruct + Mpayload = 150 + 20 = 170 kgMinitial = Minert + Mprop = 170 + 1000 = 1170 kgPropellant mass fraction η = Mprop / Minitial = 1000/1170 ≈ 0.85.
Correct Answer : 90% propellants 10% inert mass
Explanation : Much of the rocket mass is constituted by the mass of the propellants. Rest of the mass includes the rocket structure, its engines, and its payload.
Correct Answer : Cryogenic propellants need to be stored at their boiling points
Explanation : Exhaust gases leaving the ideal rocket nozzle is axially directed. Gas pressure, temperature, and velocity needs to be uniform across the nozzle axis and the gas composition remains the same along with the nozzle.
Correct Answer : Conductive heat transfer relations
Explanation : The ideal rocket engine has adiabatic walls. Hence heat transfer cannot happen through the walls. At the same time, the nozzle flow obeys isentropic (reversible and adiabatic), homogeneous expansion relations and ideal gas laws.
Correct Answer : Enthalpy
Explanation : Enthalpy comprises of flow work and internal thermal energy. Flow work arises when a fluid crosses a boundary with some velocity v. Entropy is a measure of the degree of randomness of a system.
Correct Answer : Calorically perfect gas
Explaination : A calorically perfect gas obeys ideal gas law and has constant Cp. A thermally perfect gas obeys ideal gas law, but it has a temperature dependent Cp value. A real gas is an imperfect gas that doesn’t follow the ideal gas relations and has Cp values that are temperature and pressure dependent.
Correct Answer : 21.21 cm
Explaination : Given flow over the nozzle is subsonic and incompressible. So density doesn’t vary along the nozzle and the flow velocity is less than about 0.3 times the speed of sound. In this case, the product of the nozzle area and flow velocity at any location along the nozzle length remains to be the same.A1V1 = A2 V2Given V2 = 2 V1d22/4 = (d12/4) V1/V2⇒ d2 = d1 √0.5= 21.21 cm.
Correct Answer : gas pressure
Correct Answer : Decrease in stagnation temperature
Explanation : Stagnation temperature for an isentropic process remains constant. Static temperature, static pressure, and specific volume may vary.
Correct Answer : 250 K
Explaination : In the large chamber, the flow can be assumed to be stagnant. Given that the flow is isentropic, adiabatic relations can be applied to the flow. Using P/P0 = (T/T0)(η/η-1) and taking the value of to be 1.4 (for air),T = (3/5)0.4/1.4 x 290≅250 K.
Correct Answer : Working fluid ceases to be a gas
Explanation : The maximum theoretical exhaust velocity is finite because it is limited by the thermal energy content of the fluid. If the exhaust velocity tends to a very large value, then the thermal energy content (and hence the temperature) of the working fluid rapidly decreases and may lead to the species to fall under their liquefaction or freezing point. When that happens, the working fluid changes its phase from the gaseous state.
Correct Answer : Infinite
Explaination : Pressure ratio Prc = P1 / P2, where P2 is the jet exit pressure and P1 is the chamber pressure. For the maximum theoretical value of nozzle outlet velocity, the flow has to be exhausted to vacuum. So P2 = 0 or in other words, Prc → ∞.
Correct Answer : 1000 psia, 1 atm
Explaination : Specific impulse and other design parameters of various rocket engines are compared with one another while assuming some general standard values for the chamber pressure and exit pressure. A chamber pressure of 1000 psia and an exit pressure of 1 atm are typically in use today. (1 psia = 6894.76 Pa and 1 atm = 101.325 Pa).
Correct Answer : Increase in the chamber pressure
Correct Answer : The ratio of nozzle exit area to the throat area
Explanation : Nozzle area expansion ratio is the ratio of the nozzle exit area to the nozzle’s throat area. The throat area is the smallest area and lies in between the convergent and divergent portions of a de Laval nozzle.
Correct Answer : False
Explanation : For a de Laval Nozzle, the area is inversely proportional to the ratio v/V. This follows from the continuity equation.
Correct Answer : Local conversion of kinetic to thermal energy
Explanation : In the presence of an obstruction, a stagnation zone develops locally, leading to the conversion of useful kinetic energy to thermal energy. It decreases the overall performance of the engine. In the stagnation zone, the local pressure is the stagnation pressure whose value is larger than the static pressure. Point of obstruction may or may not have normal shock waves and it depends on the flow velocity, the shape of the obstruction, nature of the gas flowing through the nozzle, etc.
Correct Answer : Minimize flow velocity
Explanation : Smooth wall surface allows smooth passage of flow through the nozzle. It won’t minimize the flow velocity. Since friction is dependent on surface roughness, it will be less for a smoother nozzle. Smoother surfaces will also increase the reflectivity of the material. Since reflectivity + transmissivity + absorptivity = 1, increasing reflectivity reduces the absorptivity of the material.
Correct Answer : Increase nozzle length
Explanation : Increase in nozzle length would lead to more material consumption, which means an increase in inert mass. It will also mean a larger exposed surface area of the rocket and while operating under atmospheric conditions will lead to higher drag.
Correct Answer : Smooth edges
Explanation : Gaps, holes, and protrusions obstruct the flow passage through the nozzle. An ideal rocket requires the flow to be uniform at any cross-section throughout the length of the nozzle.
Correct Answer : 0.992
Explanation : λ = 1/2 (1 + cosα), where α is the semi-divergence angle for a conical nozzle.Here, α = 20/2 = 10°∴ λ = 0.5 x ( 1 + 0.985)= 0.992.
Correct Answer : Increases
Explanation : λ = 1/2 (1 + cosα). As nozzle angle 2α decreases, cosα decreases, and λ also keeps decreasing. If the divergence angle is small for a conical nozzle, it causes most of the momentum to be axially directed and will lead to an increase in specific impulse.
λ = 1/2 (1 + cosα)
2α
λ
Correct Answer : 10.3 kN
Explaination : Total thrust = Corrected momentum thrust + Pressure thrustTtot = λTmom + Tpresλ = 1/2 (1 + cosα)α = 1/2 x 30° = 15°So, λ = 1/2 (1 + 0.966)= 0.983Tpres = (Pe – Pa)AeSince the operation is under standard sea level conditions, Pa = 101325 Pa.Pe = 110% of Pa.So Pe – Pa = 10% of Pa = 10132.5 PaAe = π de2/4 = 0.071∴ Tpres = 10132.5 x 0.071 = 719.075 NTmom = μe= 14 x 700 = 9800 NSo Ttot = (0.983 x 9800) + 719.075 ≅ 10.3 kN.
Correct Answer : Turn-back angle
Explanation : Turn-back angle is the difference between the nozzle angle at the exit and the angle at the inflection point. Between inflection point and the exit, when the gas flow is turned in the opposite direction, it leads to the formation of oblique compression waves.
Correct Answer : Supersonic flow
Explanation : Large divergence angles are allowed in bell-shaped nozzles because flow separation is suppressed by high relative pressure and large pressure gradient in the divergent portion. The working fluid rapidly expands in this region. But the flow need not be supersonic at all times.
Correct Answer : Sharply increases, then decrease
Explanation : For a bell-shaped nozzle, immediately after the throat section, the nozzle divergence angle increases rapidly and then it gradually decreases. It allows the flow to smoothly follow the contour without separation and leave the nozzle with more axially directed flow momentum (for eg. relative to conical nozzles).
Correct Answer : Momentum thrust
Explanation : Nozzle correction factor needs to be multiplied with momentum thrust term only. The flow momentum is affected by the change in divergence angle of a conical nozzle.
Correct Answer : 765.4 m/s
Explaination : V2av = (2π/A2)∫v2rdr integrating over 0 to r2.A2 = πd22/4= 0.011 m2∫v2rdr = ∫ (2r2 – 3r3)x 104 dr= {(2/3)r23 – (3/4)r24} x 104r2 = d2/2 = 0.06 mso V2av = (2π/0.011) x 1.34= 765.4 m/s.
Correct Answer : Heat transfer to the walls
Explanation : Because of heat transfer to the wall, the fluid layer closer to the wall will be cooler. But the boundary layer as a whole will have a higher temperature than the free stream flow because as the flow is slowed down in a boundary layer, it will lead to the conversion of flow kinetic energy to heat energy by viscous friction.
Correct Answer : Longer nozzles with high area ratios
Explanation : In longer nozzles with high area ratios, a relatively higher proportion of mass flow is found to be lying within the low-velocity region of the boundary layer. When that happens, the boundary layer plays a significant role in reducing the performance of the rocket engine.
Correct Answer : Lower momentum and higher thermal energy
Explanation : As the particle size increases, its mass increases as a function of the cube of the diameter, but the drag force on them increases only as a function of the square of its diameter. So larger particles do not move as fast as smaller ones and when they reach the nozzle exit, their temperature is higher because they give away less thermal energy.
Correct Answer : 0.54
Explaination : ζd = actual mass flow rate / ideal mass flow rate = ma / mi.For ideal conditions, ρ = P / RT = 0.15 x 106 / (287 x 590) = 0.886 kg/m3.mi = ρAV = ρQ = 0.886 x 0.53 = 0.47 kg/s.∴ ma = 1.15 x 0.47 = 0.54 kg/s.
Correct Answer : Higher chamber pressure
Explanation : Transient period happens during the start or end of the operation, i.e. when the process hasn’t achieved a steady state yet. In that period, the chamber pressure, average thrust and specific impulse values are lower than its steady-state values and the velocity variations are significant.
Correct Answer : Both Isp and c decreases
Explaination : Both specific impulse and characteristic velocity decrease as the particle fraction is increased. The loss of specific impulse is about 2% if the particle fraction is small (less than 6%). For larger values of particle fraction, the theory is not helpful and the losses can be as high as 10 to 20%.
Correct Answer : Static tests or flight tests of full-scale models
Explanation : Theoretical and actual (or delivered) performance values of a propulsion system may not be equal at all times. For evaluating delivered performance parameter values, static tests or flight test of full-scale models are used.
Correct Answer : Nozzle expansion ratio
Explanation : A thermally isolated gas tank discharges slowly than an isothermal one. This is because in the isothermal gas tank, in order to maintain the temperature constant, any kind of heat addition will result in an increase of tank pressure thereby resulting in a faster ejection.
Correct Answer : 5.09
Explaination : Thrust coefficient CF = T/PoA*.A* = πd2/4 = 78.53 cm2.∴ CF = 20,000/(500 x 1000 x 78.53 x 10-4)= 5.09.
Correct Answer : 45 mm
Explaination : Ae/A* = 9.So de2/d*2 = 9 (Since A = πd2/4).Which means de = d* x 3 = 45 mm.
Correct Answer : The vehicle tries to rotate in flight
Explanation : When the thrust line doesn’t intersect the flying vehicle’s center of mass, turning moments are generated about the COM and it tends to rotate the vehicle in flight. Moment M = r x F, where F is the thrust force and r is the position vector measure from center of mass to any point on the force vector.
Correct Answer : Turning moments
Explanation : By varying the direction of the thrust line, turning moments can be generated. Controlling the turning moments will help in attaining the desired vehicle attitude.
Correct Answer : Altering the nozzle exit area by using nozzle extensions
Explanation : Altering the nozzle exit area will not produce a thrust component not passing through the center of mass of the rocket. In that case, turning moment will not be generated.
Correct Answer : thrust axis
Explanation : Both the thrust axis and the geometric axis of the rocket nozzle exit surface are taken to be the same. For the high-thrust booster system, even a small misalignment can lead to the rocket deviating from its path during the vehicle operation period.
Correct Answer : Nozzle extensions
Explanation : Nozzle extensions are a means of improving the efficiency of rocket engines by increasing nozzle expansion ratio. It is not an irregularity in nozzle geometry. Protuberances refer to obstructions or protrusions from the nozzle wall. Out-of-round means having an unbalanced spherical or circular form or density.
Correct Answer : scarfed nozzle
Explanation : Scarfed nozzle is a passive way of thrust vector adjustment. Even though the flow Mach number of scarfed nozzles is lower compared to conical nozzles, it is more preferable because unwanted turning moments due to vehicle configuration can be avoided using it.
Explanation : Thrust coefficient, specific impulse, and thrust are independent of ambient pressure for a scarfed nozzle. Scarfed nozzle generally lies on the missile skin. The pressure forces acting on the scarfed nozzle are therefore perpendicular to the vehicle’s longitudinal axis and hence doesn’t contribute to thrust forces. Similar argumentation can be made for thrust coefficient and specific impulse as well.
Correct Answer : decreases
Explanation : With the increase in scarf angle, the motor axial specific impulse will decrease. This is because the component of exhaust velocity directed along the vehicle’s longitudinal axis decreases with the scarf angle increase.
Correct Answer : thrust coefficient
Explanation : Thrust is a weak function of thrust coefficient CF. It is in turn dependent on specific heat ratio, pressure ratio, expansion ratio, and altitude.
Correct Answer : throttle valves
Explanation : Throttle valve helps in controlling the mass flow rate thereby reducing the thrust. It usually consists of a flow controlling unit and an electro-mechanical-actuator.
Correct Answer : extremely short
Explanation : The engine might be operating at very high thrust levels. In order to bring changes to the vehicle orientation and for maneuvering purposes, time constraints might be tight and required impulses should be delivered with an extremely short pulse width.
Correct Answer : Neither mechanical and fluidic
Explanation : The force applied for vectoring the thrust is modulated in its magnitude and direction using thrust vector control systems, which generally use heavy actuators. The actuation system can be either mechanical or fluidic.
Correct Answer : 9.7°
Explaination : Required angle α = tan-1((re-rt)/L), where rt is the throat radius and re is the exit cross-section radius.∴ α = tan-1(1/2(0.02-0.007)/0.038) = 9.7°.
Correct Answer : Deploying control surfaces like fins
Explanation : Thrust modulation can be achieved by varying mass flow rate or axial component of gas flow velocity or both. Deploying control surfaces may help in controlling the attitude of the vehicle in flight within the atmosphere, but it doesn’t help in varying thrust manually.
Correct Answer : Shock Thrust Modulation
Explanation : Shock Thrust Modulation (STM) involves an axisymmetric injection of secondary flow into the diverging section of the supersonic nozzle and the creation of a normal shock by the interaction of oblique shocks. Across a normal shock, a supersonic flow becomes subsonic, thereby reducing exhaust velocity and total thrust. Fluidic Thrust Modulation is divided into STM and TSTM.
Correct Answer : Throat Shifting Thrust Modulation
Explanation : In Throat Shifting Thrust Modulation (TSTM), by shifting the throat of the nozzle, thrust is controlled. For modulation of thrust, the jet area is controlled using symmetric injection near the throat.
Correct Answer : Both assertion and reason are correct
Explaination : Usually solid propellant combustion gases are fuel rich and addition of oxygen into this control flow results in a secondary combustion. It increases both stagnation temperature and pressure of the flow. Because of pressure sensitivity on burning rate, higher stagnation pressure results in a higher mass flow rate, while the increase in stagnation temperature results in an increase in the characteristic velocity of the exit jet.
