Engine System Calibration

Although an engine has been designed to deliver a specific performance (F, 7j, m, r), a newly manufactured engine will not usually perform precisely at these nominal parameters. If the deviation from the nominal performance values is more than a few percent, the vehicle will probably not complete its intended flight course. There are several reasons for these deviations. Because of unavoidable dimensional tolerances on the hardware, the flow-pressure profile or the injector impingement (combustion efficiency) will deviate slightly from the nominal design value. Even a small change in mixture

TABLE 10-7. Typical Functions to Be Performed by Digital Computers in Monitoring and Controlling the Operation of a Liquid Propellant Rocket Engine

1. Sample the signals from significant sensors (e.g., chamber pressure, gas and hardware temperatures, tank pressure, valve position, etc.) at frequent intervals, say once, 10, 100, or 1000 times per second. For parameters that change slowly, e.g., the temperature of the control box, sampling every second or every five seconds may be adequate, but chamber pressure would be sampled at a high frequency.

2. Keep a record of all the significant signals received and all the signals generated by the computer and sent out as commands or information.

3. Control the steps and sequence of the engine start. Figure 10-12 and Table 10-6 list typical steps that have to be taken, but do not list the measured parameters that will confirm that the commanded step was implemented. For example, if the igniter is activated, a signal change from a properly located temperature sensor or a radiation sensor could verify that the ignition had indeed happened.

4. Control the shutdown of the engine. For each of the steps listed at the bottom of Table 10-6 or in Fig. 10-12 there often has to be a sensing of a pressure change or other parameter change to verify that the commanded shutdown step was taken. An emergency shutdown may be commanded by the controller, when it senses certain kinds of malfunctions, that allow the engine to be shut down safely before a dramatic failure occurs. This emergency shutdown procedure must be done quickly and safely and may be different from a normal shutdown, and must avoid creating a new hazardous condition.

5. Limit the duration of full thrust operation. For example, cutoff is to be initiated just before the vehicle attains the desired mission flight velocity.

6. Safety monitoring and control. Detect combustion instability, over-temperatures in precombustors, gas generators, or turbopump bearings, violent turbopump vibration, turbopump overspeed or other parameter known to cause rapid and drastic component malfunction, that can quickly lead to engine failure. Usually, more than one sensor signal will show such a malfunction. If detected by several sensors, the computer may identify it as a possible failure whose in-flight remedy is well known (and preprogrammed into the computer); then a corrective action or a safe shutdown may be automatically commanded by the control computer.

7. Control propellant tank pressurization. The tank pressure value has to be within an allowable range during engine operation and also during a coasting flight period prior to a restart. Sensing the activation of relief valves on the tank confirms overpressure. Automatically, the computer can then command stopping or reducing the flow of pressurant.

8. Perform automatic closed-loop control of thrust and propellant utilization (described before).

9. Transmit signals to a flying vehicle's telemetering system, which in turn can send them to a ground station, thus providing information on the engine status, particularly during experimental or initial flights.

10. Self-test the computer and software.

11. Analyze key sensor signals for deviation from nominal performance before, during, and after engine operation. Determine whether sensed quantities are outside of predicted limits. If appropriate and feasible, if more than one sensor indicates a possible out-of-limit value, and if the cause and remedy can be predicted (preprogrammed), then the computer can automatically initiate a compensating action.

ratio will cause a significant increase of residual, unused propellant. Also, minor changes in propellant composition or storage temperature (which affects density and viscosity) can cause deviations. Regulator setting tolerances or changes in flight acceleration (which affects static head) are other factors. An engine calibration is the process of adjusting some of its internal parameters so that it will deliver the intended performance within the allowed tolerance bands.

Hydraulic and pneumatic components (valves, pipes, expansion joints) can readily be water flow tested on flow benches and corrected for pressure drops and density (and sometimes also viscosity) to determine their pressure drop at rated flow. Components that operate at elevated temperatures (thrust chambers, turbines, preburners, etc.) have to be hot fired and cryogenic components (pumps, some valves) often have to be tested at the cryogenic propellant temperature. The engine characteristics can be estimated by adding together the corrected values of pressure drops at the desired mass flow. Furthermore, the ratio of the rated flows m0/rhf has to equal the desired mixture ratio r. This is shown in the example below. The adjustments include adding pressure drops with judiciously placed orifices, or changing valve positions or regulator setting.

In most pressurized feed systems the pressurizing gas is supplied from its high pressure tank through a regulator to pressurize both the fuel and the oxidizer in their respective tanks. The pressure drop equations for the oxidizer and the fuel (subscripts o and /) are given below for a pressurized feed system at nominal flows.

/>gas - (A/?gas)/ =Pl + &Pf + (&Pmj)f + (APj)f + \pfVf + Lapf (10-17)

Pgas - (A/?gas)0 =Pi + Ap0 + (ApiBj)0 + \p0v\ + Lap0 (10-18)

The gas pressure in the propellant tank is the regulated pressure pm, diminished by the pressure losses in the gasline Apm. The static head of the liquid Lap (L is the distance of the liquid level above the thrust chamber, a is the flight acceleration, and p is the propellant density) augments the gas pressure. It has to equal the chamber pressure px plus all the pressure drops in the liquid piping or valves Ap, the injector Athe cooling jacket Apj, and the dynamic flow head ¡¡pv2. If the required liquid pressures do not equal the gas pressure in the propellant tank at the nominal propellant flow, then an additional pressure drop (calibration orifice) has to be inserted. A good design provides an extra pressure drop margin for this purpose.

Two methods are available for precise control of the engine performance parameters. One uses an automatic system with feedback and a digital computer to control the deviations in real time, while the other relies on an initial static calibration of the engine system. The latter appoach is simpler and is sometimes preferred, and is still quite accurate.

The pressure balance is the process of balancing the available pressure supplied to the engine (by pumps and/or pressurized tanks) against the pressure drops plus the chamber pressure. It is necessary to do this balancing in order to calibrate the engine, so it will operate at the desired flows and mixture ratio. Figure 10-14 shows the pressure balance for one of the two branches of pro-pellant systems in a bipropellant engine with a pressurized feed system. It plots the pressure drops (for injector, cooling passages, pressurizing gas passages, valves, propellant feed lines, etc.) and the chamber pressure against the pro-pellant flow, using actual component pressure drop measurements (or estimated data) and correcting them for different flows. The curves are generally plotted in terms of head loss and volumetric flow to eliminate the fluid density as an explicit variable for a particular regulated pressure. The regulated pressure is the same for the fuel and oxidizer pressure balance and it also can be adjusted. This balance of head and flow must be made for both the fuel and oxidizer systems, because the ratio of their flows establishes the actual mixture ratio and the sum of their flows establishes the thrust. The pressure balance between available and required tank pressure, both at the desired flow, is achieved by adding a calibration orifice into one of the lines, as can be seen in Fig. 10-14. Not shown in the figure is the static head provided by the elevation of the liquid level, since it is small for many space launch systems. However, with high acceleration and dense propellants, it can be a significant addition to the available head.

For a pumped feed system of a bipropellant engine, Fig. 10-15 shows a balance diagram for one branch of the two propellants systems. The pump speed is an additional variable. The calibration procedure is usually more complex for a turbopump system, because the pump calibration curves (flow-head-power relation) can not readily be estimated without good test data and cannot easily be approximated by simple analytical relations. The flow of the propellants to a gas generator or preburner also needs to be calibrated. In this case the turbine shaft torque has to equal the torque required by the pumps and the energy losses in bearings, seals or windage. Thus a power balance must be achieved in addition to the matching of pressures and the individual propellant flows. Since these parameters are interdependent, the determination of the calibration adjustments may not always be simple. Many rocket organizations have developed computer programs to carry out this balancing.

Example 10-3. The following component data and design requirements are given for a pressurized liquid propellant rocket system similar to that in Figs. 1-3 and 10-14: fuel, 75% ethyl alcohol; oxidizer, liquid oxygen; desired mixture ratio, 1.30; desired thrust, 5000 lbf at sea level. For this propellant combustion gas k = 1.22.

Component test data: Pressure losses in gas systems were found to be negligible. Fuel valve and line losses were 9.15 psi at a flow of 9.63 lbm/sec of water. Oxidizer valve and line losses were 14.2 psi at a flow of 12.8 lbm/sec of liquid oxygen. Fuel cooling jacket prssure loss was 52 psi at a flow of 9.61 lbm/sec of water. Oxidizer side injector pressure

FIGURE 10-14. Simplified flow diagram and balance curves for the fuel or the oxidizer of a typical gas-pressurized bipropellant feed system. This diagram is also the same for a monopropellant feed system, except that it has no calibration orifice; it is calibrated by setting the proper regulated pressure.

FIGURE 10-14. Simplified flow diagram and balance curves for the fuel or the oxidizer of a typical gas-pressurized bipropellant feed system. This diagram is also the same for a monopropellant feed system, except that it has no calibration orifice; it is calibrated by setting the proper regulated pressure.

Rated H

Available pump pressure (characteristic pump curves)

Control valve adjustment drop

Available pump pressure (characteristic pump curves)

Rated H

Control valve adjustment drop

Valve and ; line losses Injection drop

Rated Q Pump capacity or flow

FIGURE 10-15. Simplified diagram of the balance of available and required feed pressures versus flow for one of the propellants in a rocket engine with a turbopump feed system. Chamber pressure is increased by liquid column.

Rated Q Pump capacity or flow

Valve and ; line losses Injection drop

FIGURE 10-15. Simplified diagram of the balance of available and required feed pressures versus flow for one of the propellants in a rocket engine with a turbopump feed system. Chamber pressure is increased by liquid column.

drop was 90.0 psi at 10.2 lb/sec of oxygen flow under thrust chamber operating conditions. Fuel side injector pressure drop was 48.3 psi at 10.2 lb/sec of fuel flow under thrust chamber operating conditions. Average results of several sea-level thrust chamber tests were: thrust = 5410 lbf; mixture ratio = 1.29; specific impulse = 222 sec; chamber pressure = 328 psia; nozzle area ratio = 4.0. Determine regulator setting and size and location of calibration orifices.

SOLUTION. First, the corrections necessary to obtain the desired thrust chamber conditions have to be determined. The experimental thrust chamber data must be adjusted for deviations in mixture ratio, thrust, and specific impulse. The variation of specific impulse with mixture ratio is determined from experimental data or (on a relative basis) from theoretical calculations similar to those that are the basis of Fig. 5-1. Because the value of Is at the desired mixture ratio of 1.30 is within 0.08% of the value of Is under the actual test conditions (r = 1.29), any mixture ratio correction of Is is neglected here.

The correction of the specific impulse for chamber pressure is made next. The specific impulse is essentially proportional to the thrust coefficients as determined from Eq. 3-30. For k = 1.22, and the pressure ratios pjp^ = 328/14.7 = 22.2 and 300/14.7= 20.4, the values of CF can be calculated as 1.420 and 1.405, respectively. In this calculation p2 has to be determined for isentropic conditions, such as those in Figs. 3-7 or 3-8 for the given nozzle area ratio. The sea-level specific impulse is therefore corrected to Is = 222 (1.405/1.420)= 220 sec. The chamber pressure has to be reduced from 328 psi to a lower value in order to bring the thrust from its test value of 5410 lbf to the design value of 5000 lbf. In accordance with Eq. 3-31, F = CFA,px. The chamber pressure is inversely proportional to the thrust coefficient CF and proportional to the thrust, and therefore

Pl/P\ = (F{/F[)(C'f/Cf) The primes refer to the component test condition.

The desired total propellant flow is, from Eq. 2-5, w = F/Is = 5000/220 = 22.7 lbf/sec

For a mixture ratio of 1.3, the desired fuel and oxidizer flows are obtained from Eqs. 6-3 and 6-4 as wy = 9.9 lbf/sec and w0 = 12.8 lbf/sec. Next, the various component pressure drops are corrected to the desired flow values and to the corrected propellant densities in accordance with Eq. 8-2, which applies to all hydraulic devices. By neglecting variations in discharge coefficients, this equation can be rewritten into a convenient form:

With this equation and the specific gravity values (from Fig. 7-1) of 1.14 for oxygen, 0.85 for diluted ethyl alcohol, and 1.0 for water, the new pressure drops for the corrected flow conditions can be found, and these are tabulated below with flow values given in pounds per second and pressure values in pounds per square inch.

Component Test Data Design Conditions

Component

Fluid

w

A p

Fluid

w

A p

Fuel injector

Fuel

Was this article helpful?

0 0
Project Management Made Easy

Project Management Made Easy

What you need to know about… Project Management Made Easy! Project management consists of more than just a large building project and can encompass small projects as well. No matter what the size of your project, you need to have some sort of project management. How you manage your project has everything to do with its outcome.

Get My Free Ebook


Post a comment